Aerial propulsion system with improved propulsion efficiency

Abstract

An aviation propulsion system (1) in which the average inner radius of a high-pressure turbine (7) is at least equal to formula (I): # imgabs0 # wherein: R average int is the average inner radius of the high-pressure turbine (7) in millimeters (mm); d is the diameter of the fan rotor (9), and the unit is millimeter (mm); bPR is the bypass ratio of the propulsion system (1); te is the inlet temperature of the first turbine (8) when the propulsion system (1) is static and at a take-off rated value under standard atmospheric pressure and sea level conditions, and is expressed by DEG C (DEG C); the Tref is equal to 273 K; gamma is the adiabatic coefficient of air; and A = 593 (DEG C)-1 / 2 and B = 6.5 mm (mm).

Description

Technical Field

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

[0003] Technical research has significantly improved the environmental performance of aircraft. Applicants have considered factors affecting all stages of design and development 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 for an increase in the inlet pressure of the combustion chamber gases, thus further improving the overall efficiency of the propulsion system. Consequently, increasing the overall compression ratio requires increasing the compression ratio of the high-pressure compressor and / or the low-pressure compressor, particularly since, for the reasons mentioned above, the fan compression ratio is also being sought to be reduced. One consequence is a smaller high-pressure turbine, which is subject to greater mechanical loads, particularly at the blade roots. Summary of the Invention

[0006] One object of the present application is to optimize a propulsion system in order to increase the efficiency of the propulsion system without overloading and / or thermally overloading the high-pressure turbomachinery.

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

[0008] a first turbine configured to drive the first compressor via a first shaft;

[0009] a second turbine configured to drive the second compressor via a second shaft configured to rotate at a higher speed than the first shaft;

[0010] - a fan rotor (9), which is connected to the fan shaft;

[0011] a 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; wherein the average inner radius of the second turbine is at least equal to:

[0012]

[0013] Where: R 平均 _int is the average inner radius of the second turbine, in millimeters;

[0014] 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 blades of the fan rotor;

[0015] BPR is the propulsion system bypass ratio and is measured when the propulsion system is stationary and at takeoff rating under standard atmospheric pressure and sea level conditions;

[0016] T e is the inlet temperature of the first turbine when the propulsion system is stationary and at takeoff rating under standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C);

[0017] T ref =273K;

[0018] GAMMA is the adiabatic coefficient of air; and

[0019] A=593(℃) -1 / 2 And B=6.5 millimeters (mm).

[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 average inner radius of the second impeller is at most equal to 300 mm;

[0022] - a hub-to-tip ratio of the second turbine is greater than 0.77 and less than 0.90;

[0023] - the propulsion system further comprises an inter-turbine housing mounted on the bearing assembly, the inter-turbine housing extending between the first turbine and the second turbine;

[0024] - the inter-turbine housing comprises a series of guide vanes configured to straighten the air flow at the inlet of the first turbine;

[0025] - the propulsion system includes between twenty and thirty guide vanes;

[0026] -The second turbine is a two-stage turbine;

[0027] - the second compressor comprises at least eight stages and at most eleven stages;

[0028] The fan rotor has a diameter between 80 inches and 185 inches, inclusive, preferably between 85 inches and 120 inches, inclusive, such as approximately 90 inches;

[0029] The fan rotor is ducted, and the bypass ratio of the propulsion system is greater than or equal to 10, for example, between 10 and 35 inclusive, preferably between 10 and 18 inclusive;

[0030] - the fan rotor is unducted, and the propulsion system has a bypass ratio greater than or equal to 40, such as between 40 and 80, inclusive;

[0031] - a reduction ratio of the reduction mechanism greater than or equal to 2.5, preferably greater than or equal to 3.0 and less than or equal to 11.0;

[0032] - the first turbine comprises at least three stages and at most five stages; and / or

[0033] - The first compressor comprises at least two stages and at most four stages.

[0034] According to a second aspect, the present application proposes an aircraft comprising at least one propulsion system according to the first aspect, the propulsion system being fixed to the aircraft via a pylon.

[0035] According to a third aspect, the present application provides a method for determining the size of a propulsion system or manufacturing a propulsion system, the propulsion system comprising a reduction mechanism and a second turbine, the reduction mechanism coupling 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 being configured to rotate at a speed higher than that of the first turbine, the size of the second turbine being determined such that an average inner radius of the second turbine is at least equal to:

[0036]

[0037] Where: R平均 _int is the average inner radius of the second turbine, in millimeters;

[0038] 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 blades of the fan rotor;

[0039] BPR is the propulsion system bypass ratio and is measured when the propulsion system is stationary and at takeoff rating under standard atmospheric pressure and sea level conditions;

[0040] T e is the inlet temperature of the first turbine when the propulsion system is stationary and at takeoff rating under standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C);

[0041] T ref =273K;

[0042] Gamma is the adiabatic coefficient of air; and

[0043] A=593(℃) -1 / 2 And B=6.5 millimeters (mm). BRIEF DESCRIPTION OF THE DRAWINGS

[0044] Other features, objects and advantages of the present invention 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:

[0045] 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;

[0046] 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;

[0047] Figure 3 is a schematic cross-sectional view of an example of a star-type reduction mechanism;

[0048] Figure 4 is a schematic cross-sectional view of an example of a planetary reduction mechanism;

[0049] Figure 5 is an example of an aircraft that may include at least one propulsion system according to the first or second embodiment; and

[0050] Figure 6 is a flow chart illustrating an example of steps in a sizing or manufacturing method according to one embodiment.

[0051] In all figures, similar elements have the same reference numerals. DETAILED DESCRIPTION

[0052] 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. 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).

[0053] 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, with impellers (rotors) 7a and 8a rotating behind them.

[0054] 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), 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.

[0055] 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.

[0056] 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.

[0057] 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 passes through the turbine sections 7, 8, which receive energy from the combustion chamber 6, which causes the rotation of the rotors of the turbine sections 7, 8. The rotation of the rotors of the turbine sections 7, 8 in turn drives the rotors of the compressor sections 4, 5 and the rotor portion 9 of the fan section 2 to rotate.

[0058] In a two-shaft 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 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-shaft 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.

[0059] 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 appropriate, 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.

[0060] 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 the stator part of the propulsion system 1. 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.

[0061] The fan section 2 may also include a fan stator 16 or straightener, which includes blades 17 mounted on a hub 18 of the fan stator 16 and has 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 have a variable setting. In a manner similar to the rotor blades 14, the roots of the stator blades 17 may be pivotally mounted along a setting axis X and connected to a pitch changing mechanism 15a, which is generally different from the pitch changing mechanism 15 of the fan rotor 9, and the setting is adjusted by the pitch changing mechanism 15a according to the flight phase.

[0062] 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. A high bypass ratio is defined herein as 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 stationary, not installed, and at takeoff ratings under standard atmospheric pressure (as defined by the 3rd edition of the International Civil Aviation Organization (OACI) Manual, Document 7488 / 3) and sea level conditions (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 means that the measurements are taken when the propulsion system 1 is in a test bench (and not installed on the aircraft 100), which is easier to perform. 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.

[0063] 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.

[0064] 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 is regulated to the first order by the propulsion efficiency, which is beneficially influenced by minimizing the change 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 made up 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 during its passage through the fan section 2. Therefore, the propulsion 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 propulsion efficiency. To optimize the propulsion efficiency of the propulsion system 1, the fan pressure ratio (corresponding to the ratio between the average pressure at the outlet of the fan stator 17 (or, in the absence of a stator, the fan rotor 9) and the average pressure at the inlet of the fan rotor 9) is less than or equal to 1.70, preferably less than or equal to 1.50, and 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 ).

[0065] The propulsion system 1 is configured to provide a thrust between 18,000 lbf (80,068 N) and 51,000 lbf (222,411 N), preferably between 20,000 lbf (88,964 N) and 35,000 lbf (155,688 N).

[0066] The fan section 2 may be ducted or non-ducted. 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 housed.

[0067] The ducted fan section 2 includes a fan rotor 9 extending upstream of the fan stator. The blades 17 of the fan stator are generally referred to as outlet guide vanes (OGV) and have a fixed setting relative to the hub of the fan stator. In addition, preferably, the bypass ratio of the propulsion system 1 is greater than or equal to 10, for example, between 10 and 35 (including 10 and 35), preferably between 10 and 18 (including 10 and 18). 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. Then, the pressure ratio of the fan can be between 1.20 and 1.45.

[0068] 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 settings. A propulsion system comprising at least one unducted fan rotor 9 is also 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 "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 arranged at the front of the main body 3 to form a pull-type structure. As a variant, 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 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 straightener 16 are also variably settable blades.

[0069] The absence of a fairing surrounding the fan section 2 makes it possible to significantly increase the bypass ratio, without the propulsion system 1 being adversely affected by the mass of the casing or nacelle intended 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 of the one or more fan rotors 9 may also be between 210 m / s and 260 m / s. Preferably, the pressure ratio of the fan may be between 1.05 and 1.20.

[0070] The reduction mechanism 19 may include, for example, a reduction mechanism 19 having a planetary gear train, such as a single-stage or two-stage reduction mechanism 19 of a "planetary" or "star" type. According to a first variation, the reduction mechanism 19 may be a star-type ( Figure 3), and includes 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 to rotate 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 meshing 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 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 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 casing of the compressor sections 4, 5).

[0071] 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 .

[0072] 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 about 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.

[0073] The limit speed of the low-pressure shaft 11 is between 8500 rpm and 12000 rpm, preferably between 9000 rpm and 11000 rpm, which corresponds to the absolute maximum speed that the low-pressure shaft 11 may encounter during the entire flight (according to the European certification regulation EASA CS-E 740 (or according to the US certification regulation 14-CFR Part 33)). The limit speed corresponds to the maximum rotational speed when the propulsion system is intact (and possibly at the end of its life). Therefore, the low-pressure shaft 11 is likely to reach the limit speed under flight conditions. This limit speed forms part of the data stated in the engine certification (type certificate data sheet). In practice, this rotational speed is usually used as a reference speed for determining the size of the propulsion system 1 and in some certification tests (such as blade loss or rotor integrity tests).

[0074] In order to improve the efficiency of the propulsion system 1 , the dimensions of the high-pressure turbine 7 are determined so that the average inner radius R of the high-pressure turbine 7 is 平均 _int is at least equal to:

[0075]

[0076] Where: R 平均 _int is expressed in millimeters (mm);

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

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

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

[0080] T ref =273K;

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

[0082] A=593(℃) -1 / 2 And B=6.5 millimeters (mm).

[0083] As previously described, all of these parameters, except for the mean inside radius and diameter, which are determined when the propulsion system 1 is cold, are determined when the propulsion system 1 is stationary, uninstalled, and at takeoff ratings at standard atmospheric pressure (as defined by the 3rd edition of the International Civil Aviation Organization (OACI) Manual, Document 7488 / 3) and sea level conditions.

[0084] The average inner radius R of the high-pressure turbine 平均 _int is equal to the arithmetic mean of the inner radius R1 of the rotor 7b (blade wheel) of the high-pressure turbine 7. At a given stage, the inner radius R1 of the rotor 7b corresponds to the distance between the outer radial surface of the hub of the rotor 7b (the outer radial surface radially internally delimiting the flow path in the rotor 7b) and the axis of rotation X of the high-pressure turbine 7, in a plane perpendicular to the axis of rotation X of the propulsion system 1 when the propulsion system 1 is stationary, and located halfway between the leading edge 7c and the trailing edge 7d of the blade at the hub of the rotor 7b (at 50% of the chord at the blade root).

[0085] The dimensions of the high-pressure turbine 7 are determined to comply with the formula (1) defined above, so that a compromise can be obtained between an acceptable aerodynamic load at the blade root and sufficient expansion work to increase the compression ratio of the high-pressure compressor 5 (for the same number of stages), thereby increasing the overall compression ratio of the propulsion system 1. Alternatively, the dimensions of the high-pressure turbine 7 are determined so that the compression ratio of the low-pressure compressor 4 can be reduced (for the same overall compression ratio), and therefore the average radius of the low-pressure compressor 4 for the same load can be reduced, which makes it possible to improve the bypass ratio BPR of the propulsion system 1. The high-pressure turbine 7 can indeed rotate at a sufficiently high speed (in revolutions per minute) to enable the high-pressure compressor 5 to achieve a compression ratio greater than 21 (in takeoff rating) without increasing the number of stages in the high-pressure compressor 5. Thus, for a two-stage high-pressure turbine 7, the high-pressure compressor 5 can include at least eight stages and at most eleven stages, for example nine stages.

[0086] The diameter of the high-pressure turbine 7 can also be large enough to enable the passage and, if appropriate, radial displacement of the supercritical low-pressure shaft 11 .

[0087] In one embodiment, the average inner radius R of the high pressure turbine 11 is 平均 _int is at most equal to 300 mm. In a propulsion system 1 with a high bypass ratio, formula (1) also gives an average inner radius R of at least 200 mm. 平均 _int.

[0088] The inlet temperature of the low-pressure turbine is T e It can be between 950°C and 1230°C.

[0089] In order to obtain the average inner radius R 平均 _intThe high-pressure turbine 7 that complies with formula (1), in particular, can be constructed with respect to the cross-section of the flow path at the distributor at the outlet of the combustion chamber 6 and at the inlet of the high-pressure turbine 7 so that, at the takeoff rating (as described above), the Mach number in the path cross-section is approximately equal to 1. To this end, the average radius of the high-pressure turbine 7 is determined by the acceptable aerodynamic load in the high-pressure turbine 7, which itself is defined by the energy that must be supplied to the high-pressure compressor 5 and the rotational speed of the high-pressure shaft 10 in order to obtain the required compression ratio. The cross-section of the flow path in the high-pressure turbine 7 depends on the reduced flow in the flow path at a given Mach number. Here, the size of the cross-section of the flow path can therefore be determined so that the Mach number in the path cross-section is equal to 1. Average inner radius R 平均 _int is derived from the cross section of the path dimensioned in this way and the mean radius determined from the aerodynamic loads.

[0090] The rotation speed of the high-pressure turbine 7 may be between 15,000 rpm and 27,000 rpm.

[0091] The high-pressure turbine 7 conforming to 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 7b (blade wheel) of the high-pressure turbine 7, the hub-to-tip ratio being 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, the plane being located at 50% of the chord at the root of the blades of the rotor 7b of the most downstream portion of the high-pressure turbine 7 (i.e., the last stage of the high-pressure turbine 7). When the propulsion system 1 is stationary, the outer radius R2 of the high-pressure turbine 7 corresponds to the distance between the tip 7e of the blade of the rotor 7b 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 is as defined above.

[0092] This high-pressure turbine 7 has an optimized outlet cross section S s . In 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 of between 0.77 and 0.90 is combined with the optimized inlet cross-section Se and rotational speed as described above, making it possible not only to obtain a more efficient high-pressure turbine 7 in a suitable occupied space (due to the adapted rotational speed of the high-pressure turbine 7), but also to optimize the outlet surface for expanding the gas at the outlet of the combustion chamber 5. Therefore, the high-pressure turbine 7 is dimensioned to obtain a hub-to-tip ratio of between 0.77 and 0.90, which makes the high-pressure turbine 7 more efficient and therefore reduces the specific consumption of the propulsion system 1 without compromising the mechanical load or thermal load of the propulsion system 1.

[0093] Including the average inner radius R 平均 _intThe propulsion system 1 of the high-pressure turbine 7 that complies with formula (1) can have a total 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 total 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.

[0094] 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.

[0095] 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.

[0096] Comparison example:

[0097] (a) Propulsion systems with ducted fans:

[0098] The engine 1 is a twin-shaft propulsion system including a ducted fan section corresponding to the current technical standards (at the filing date of the present application) for which improvement is sought.

[0099] The engine 2 is a dual-shaft propulsion system 1 including a ducted fan consistent with the teachings of the present application.

[0100]

[0101]

[0102] Since the average inner radius of the high-pressure turbine of engine 1 is less than 236.9 mm, it does not comply with the claimed formula. In contrast, the average inner radius of the high-pressure turbine 7 of engine 2 complies with the claimed formula. It can be seen that for the same number of stages, the aerodynamic load N2 at the blade root of engine 2 is 2 S is lower than the aerodynamic load of the engine 1 while maintaining sufficient expansion work to maintain the compression ratio of the high pressure compressor.

[0103] In order to move from (reference) engine 1 to engine 2 (in accordance with the present disclosure), the overall compression 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. In addition, the fan diameter D and the bypass ratio BPR are increased, which makes it possible to improve the propulsion efficiency of engine 2.

[0104] (b) Propulsion systems with unducted fans:

[0105] The engine 3 is a two-shaft propulsion system including a non-ducted fan section corresponding to the current technical standards (at the filing date of the present application) for which improvement is sought.

[0106] Engine 4 is a dual-shaft propulsion system 1 including an unducted fan consistent with the teachings of the present application.

[0107]

[0108]

[0109] Since the average inner radius of the high-pressure turbine of engine 3 is less than 255.9 mm, it does not comply with the claimed formula. In contrast, the average inner radius of the high-pressure turbine 7 of engine 4 complies with the claimed formula. Similar to ducted engine 2, for the same number of stages, the aerodynamic load N2 at the blade root of engine 4 is 2 S is lower than the aerodynamic load of the engine 1 while maintaining sufficient expansion work to maintain the compression ratio of the high pressure compressor.

[0110] To move from (reference) Engine 3 to Engine 4 (consistent with the present disclosure), the fan diameter D is reduced, and the bypass ratio is also reduced to facilitate the integration of Engine 4 under the wing. The fan pressure ratio is also increased very slightly (while remaining below 1.45) to maintain fan thrust. The overall compression ratio is also increased, and the inlet temperature of the high-pressure turbine is also increased to improve the thermal efficiency of the high-pressure body and compensate for the loss in specific consumption associated with the increase in the fan pressure ratio. The number of stages in the low-pressure compressor is also increased to increase the overall compression ratio.

[0111] The reduction in fan diameter allows for a more compact propulsion system, making it easier to install the propulsion system on an aircraft. The reduction in propulsion system mass is beneficial for reducing aircraft consumption.

Claims

1. An aviation propulsion system (1), comprising: a first turbine (8) configured to drive the first compressor (4) 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 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); Wherein, the average inner radius of the second turbine (7) is at least equal to: Where: R 平均 _int is the average inner radius of the second turbine (7), in millimeters (mm); D is the diameter of the fan rotor (9) in millimeters (mm) and is measured in a plane perpendicular to the axis of rotation 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 and at a takeoff rating under standard atmospheric pressure and sea level conditions; T e is the inlet temperature of the first turbine (8) when the propulsion system (1) is at rest and at takeoff rating under standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C); T ref =273K; Gamma is the adiabatic coefficient of air; and A=593(℃) -1 / 2 And B=6.5 millimeters (mm).

2. The propulsion system (1) according to claim 1, wherein The average inner radius of the second turbine (7) is at most equal to 300 mm.

3. The propulsion system (1) according to claim 1 or 2, wherein: The hub-to-tip ratio of the second turbine (7) is greater than 0.77 and less than 0.

90.

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

5. The propulsion system (1) according to claim 4, wherein: The inter-turbine housing (24) includes a series of guide vanes (25) configured to straighten the air flow at the inlet of the first turbine (8).

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

7. The propulsion system (1) according to any one of claims 1 to 6, wherein: The second turbine (7) is a two-stage turbine.

8. The propulsion system (1) according to any one of claims 1 to 7, wherein: The second compressor (5) comprises at least eight stages and at most eleven stages.

9. The propulsion system (1) according to any one of claims 1 to 8, wherein: The fan rotor (9) has a diameter between 80 inches (2032 mm) and 185 inches (4699 mm) inclusive, preferably between 85 inches (2159 mm) and 120 inches (3048 mm) inclusive, for example, approximately 90 inches (2286 mm).

10. The propulsion system (1) according to any one of claims 1 to 9, wherein: The fan rotor (9) is of ducted type, and 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, preferably between 10 and 18 and including 10 and 18.

11. The propulsion system (1) according to any one of claims 1 to 10, wherein: The fan rotor (9) is of non-ducted type, and the bypass ratio of the propulsion system (1) is greater than or equal to 40, for example, between 40 and 80, including 40 and 80.

12. The propulsion system (1) according to any one of claims 1 to 11, wherein: in, The reduction ratio of the reduction mechanism (19) is greater than or equal to 2.5, preferably greater than or equal to 3.0 and less than or equal to 11.

0.

13. The propulsion system (1) according to any one of claims 1 to 12, wherein: The first 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 first compressor (4) comprises at least two stages and at most four stages.

15. An aircraft (100) comprising at least one propulsion system (1) according to any one of claims 1 to 14, the propulsion system being fixed to the aircraft (100) via a pylon.

16. A method for determining the size of a propulsion system (1), the propulsion system comprising a reduction mechanism (19) coupling a first turbine (8) and a fan rotor (9) so as to drive the fan rotor (9) at a speed lower than the speed of the first turbine (8), and a second turbine (7) configured to rotate at a speed higher than the speed of the first turbine (8), the second turbine (7) being sized such that the average inner radius of the second turbine (7) is at least equal to: in: R 平均 _int is the average inner radius of the second turbine (7), in millimeters (mm); D is the diameter of the fan rotor (9) in millimeters (mm) and is measured in a plane perpendicular to the axis of rotation 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 and at a takeoff rating under standard atmospheric pressure and sea level conditions; T e is the inlet temperature of the first turbine (8) when the propulsion system (1) is at rest and at takeoff rating under standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C); T ref =273K; Gamma is the adiabatic coefficient of air; and A=593(℃) -1 / 2 And B=6.5 millimeters (mm).

17. A method for manufacturing a propulsion system (1), comprising the following steps: - dimensioning of the propulsion system according to claim 16; and - Manufacturing the propulsion system.

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