Aerial propulsion system with improved propulsion efficiency

Abstract

An aviation propulsion system (1) comprises a high-pressure compressor (5) having an average outer radius satisfying formula (I) # imgabs0 # wherein: Rext5 is the average outer radius of the high-pressure compressor (5) in millimeter (mm), Rext2 is the average outer radius of the high-pressure compressor (5), Rext3 is the average outer radius of the high-pressure compressor (5) in millimeter (mm); d9 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 maximum temperature at the inlet of the first turbine (8) when the propulsion system (1) is stationary and in a takeoff mode, in degree Celsius (DEG C); n is the rotational speed of the second shaft (10) in rotational speed per minute (tr / min), and N is measured when the propulsion system (1) is stationary and in takeoff mode; s7 is the average surface of the second turbine (7) in square millimeters (mm2); and A = 1 (trr / min) 2 mm / DEG C and B = 8977 (trr / min) 2 * mm2.

Description

Technical Field

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

[0002] The propulsion system generally includes, from upstream to downstream along the direction of gas flow, a fan section, a compressor section, a combustor, and a turbine section. The compressor section may include a low-pressure compressor and a high-pressure compressor, and the turbine section may specifically include a high-pressure turbine and a low-pressure turbine. The high-pressure compressor is 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] 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. Summary of the Invention

[0005] One object of the present application is to optimize the performance of the propulsion system, in particular the footprint of the propulsion system and the overall dynamics of the propulsion system, without compromising the efficiency or the mechanical strength of the propulsion system.

[0006] To this end, according to a first aspect, an aviation propulsion system is proposed, comprising:

[0007] - a fan rotor, the fan rotor being connected to the fan shaft;

[0008] a first turbine configured to drive a fan rotor (9) 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 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;

[0011] The average outer radius of the second compressor satisfies the following formula:

[0012]

[0013] Where: R ext_5 is the average outer radius of the second compressor in millimeters;

[0014] D9 is the diameter of the fan rotor in millimeters, and D9 is measured in a plane perpendicular to the axis of rotation of the fan rotor at the intersection between the tip and the leading edge of the blades of the fan rotor;

[0015] 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;

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

[0017] N is the rotational speed of the second shaft and is measured with the propulsion system at rest, at takeoff rating, at standard atmospheric pressure and at sea level, in revolutions per minute (tr / min);

[0018] S7 is the average surface area of ​​the second turbine in square millimeters (mm 2 ) means; and

[0019] A=1(tr / min) 2 .mm / ℃ and B=8977(tr / min) 2 *mm 2 .

[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 outer radius of the second compressor also satisfies the following formula:

[0022]

[0023] Where: C = 7668 2 *mm 2 ;

[0024] -The average outer radius of the second compressor also satisfies the following formula:

[0025] R ext_5 >(E*n5+F)*R ext_7

[0026] Where: n5 is the number of stages of the second compressor; R ext_7 is the average outer radius of the second turbine in millimeters (mm); and E = -23.9*10 -3 Square millimeters (mm 2 ) and F = 0.96;

[0027] -The average outer radius of the second compressor also satisfies the following formula:

[0028] R ext_5 <(E*n5+G)*R ext_7

[0029] Where: n5 is the number of stages of the second compressor; R ext_7 is the average outer radius of the second turbine in millimeters (mm); and E = -23.9*10 -3 Square millimeters (mm 2 ) and G = 1.06;

[0030] - The second turbine is a two-stage turbine and the second compressor is axial;

[0031] - the second compressor comprises at least eight stages and at most eleven stages;

[0032] The fan rotor has a diameter between 2032 mm and 4699 mm inclusive, for example between 2159 mm and 3048 mm inclusive, for example approximately 2286 mm;

[0033] The fan section is of a ducted type, and the propulsion system has a bypass ratio greater than or equal to 10, such as between 10 and 35 inclusive, such as between 10 and 18 inclusive;

[0034] The fan section is unducted, and the propulsion system has a bypass ratio greater than or equal to 40, such as between 40 and 80, inclusive;

[0035] - a hub-to-tip ratio at the inlet of the second compressor between 0.41 and 0.60;

[0036] - the rotational speed of the second shaft is greater than or equal to 15,000 rpm and less than or equal to 27,000 rpm;

[0037] - 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;

[0038] - the first turbine comprises at least three stages and at most five stages;

[0039] - the first turbine also drives a first compressor via a first shaft, the first compressor comprising at least two stages and at most four stages; and / or

[0040] The fan section also has a fan pressure ratio, corresponding to the pressure ratio between the outlet of the fan rotor and the inlet of the fan rotor, which is less than or equal to 1.45, for example less than or equal to 1.30.

[0041] According to a second aspect, an aircraft is proposed, comprising at least one propulsion system according to the first aspect, the propulsion system being fixed to the aircraft via a mast.

[0042] According to a third aspect, a method for determining the size of or manufacturing a propulsion system is provided. The propulsion system includes a 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, and a second turbine configured to rotate at a speed higher than that of the first turbine, wherein an average outer radius of the second compressor satisfies the following formula:

[0043]

[0044] Where: R ext_5 is the average outer radius of the second compressor in millimeters;

[0045] D9 is the diameter of the fan rotor in millimeters and is measured in a plane perpendicular to the axis of rotation of the fan rotor at the intersection between the tip and the leading edge of the fan rotor blades. Figure 1 and Figure 2 This is a partial view, so diameter D9 is only partially visible;

[0046] BPR is the propulsion system bypass ratio and is measured when the propulsion system is stationary and at takeoff rating at standard atmospheric pressure and sea level;

[0047] T e is the maximum inlet temperature of the first turbine when the propulsion system is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C);

[0048] N is the rotational speed of the second shaft and is measured with the propulsion system at rest, at takeoff rating, at standard atmospheric pressure and at sea level, in revolutions per minute (tr / min);

[0049] S7 is the average surface area of ​​the second turbine in square millimeters (mm 2 ) means; and

[0050] A=1(tr / min) 2 .mm / ℃ and B=8977(tr / min) 2 *mm 2 .

[0051] Some preferred but non-limiting features of the sizing method or manufacturing method according to the third aspect are the following features taken alone or in combination:

[0052] -The average outer radius of the second compressor satisfies the following formula:

[0053]

[0054] Where: C = 7668 (tr / min) 2 *mm 2 ;

[0055] -The average outer radius of the second compressor also satisfies the following formula:

[0056] R ext_5 >(E*n5+F)*R ext_7

[0057] Where: n5 is the number of stages of the second compressor; R ext_7 is the average outer radius of the second turbine in millimeters (mm); and E = -23.9*10 -3 Square millimeters (mm 2 ) and F=0.96; and / or

[0058] -The average outer radius of the second compressor also satisfies the following formula:

[0059] R ext_5 <(E*n5+G)*R ext_7

[0060] Where: n5 is the number of stages of the second compressor; R ext_7 is the average outer radius of the second turbine in millimeters (mm); and E = -23.9*10 -3 Square millimeters (mm 2 ) and G=1.06.

[0061] According to a fourth aspect, a method for manufacturing an aviation propulsion system is provided, comprising the following steps:

[0062] - determining the dimensions of the aerospace propulsion system according to the dimensioning method of the third aspect; and

[0063] -Manufacturing of aviation propulsion systems. BRIEF DESCRIPTION OF THE DRAWINGS

[0064] 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:

[0065] Figure 1 is a schematic partial cross-sectional view of an example of a propulsion system according to the first embodiment, wherein the fan section is ducted;

[0066] Figure 2 is a schematic partial cross-sectional view of an example of a propulsion system according to the first embodiment, wherein the fan section is unducted;

[0067] Figure 3 is a schematic cross-sectional view of one example of a speed reduction mechanism according to a first modification;

[0068] Figure 4 is a schematic cross-sectional view of one example of a speed reduction mechanism according to a second modification;

[0069] Figure 5 is an example of an aircraft according to the first or second embodiment that may include at least one propulsion system;

[0070] Figure 6 is a flow chart illustrating an example of steps in a sizing or manufacturing method according to one embodiment.

[0071] In all figures, similar elements have the same reference numerals. DETAILED DESCRIPTION

[0072] 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).

[0073] 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.

[0074] 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.

[0075] 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.

[0076] 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.

[0077] 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. This causes the rotation of the rotors of the turbine sections 7, 8, which in turn drives the rotors of the compressor sections 4, 5 and the rotor part 9 of the fan section 2 to rotate.

[0078] In a dual-body propulsion system 1, compressor sections 4 and 5 may include a low-pressure compressor 4 and a high-pressure compressor 5. Turbine sections 7 and 8 may include a high-pressure turbine 7 and a low-pressure turbine 8. The rotor of high-pressure turbine 7 drives the rotor of high-pressure compressor 5 via a high-pressure shaft 10. The rotor of high-pressure turbine 8 drives the rotor of low-pressure compressor 4 and the rotor portion 9 of fan section 2 via a low-pressure shaft 11. Thus, main body 3 includes a high-pressure main body, which includes high-pressure compressor 5, high-pressure turbine 7, and high-pressure shaft 10, and a low-pressure main body, which includes fan section 2, low-pressure compressor 4, high-pressure turbine 8, and low-pressure shaft 11. The rotational speed of the high-pressure main body is greater than that of the low-pressure main body. In a three-shaft propulsion system 1, turbine sections 7 and 8 also include an intermediate turbine, located between high-pressure turbine 7 and high-pressure turbine 8 and configured to drive the rotor of low-pressure compressor 4 via the intermediate shaft. The fan rotor 9 and the rotor of high-pressure compressor 5 are still driven by the low-pressure shaft 11 and high-pressure shaft 10, respectively.

[0079] 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.

[0080] 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.

[0081] The fan section 2 may also include a fan stator 16 or a straightener comprising guide vanes 17 mounted on a hub 18 of the fan stator 16, the straightener having the function of straightening the secondary air flow F2 flowing at the outlet of the fan rotor 9. The vanes 17 of the fan stator 18 may be fixed relative to the hub 18 or may have a variable setting. 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.

[0082] Furthermore, the fan rotor 9 comprises at least twelve blades 14 and at most twenty-four blades 14, for example at least sixteen blades 14 and at most twenty-two blades 14. The number of blades 16 in the fan stator 17 depends on the acoustic standards defined for the propulsion system 1 and is at least equal to the number of blades 14.

[0083] 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, 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 is at rest, not installed, at takeoff rating, at standard atmospheric pressure (as defined by the 3rd edition of the International Civil Aviation Organization (OACI) Manual, Document 7488 / 3), and at sea level conditions (referred to as sea level standard (SLS) conditions). It should be noted that in this application, parameters (pressure, flow, thrust, speed, etc.) are systematically determined under these conditions. "Not installed" here refers to measurements taken when the propulsion system 1 is in a test stand (and not installed on the aircraft 100), which is easier to perform. However, when the propulsion system 1 is cooled, that is, when the propulsion system has been stopped for a sufficient time so that the components of the propulsion system are at normal temperature, distances (lengths, radii, diameters, etc.) are measured at normal temperature (about 20°C), and it should be understood that these dimensions change very little compared to the conditions when the propulsion system 1 is at takeoff rating.

[0084] 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 high-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 lower rotational speed than the high-pressure turbine 8.

[0085] 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 high-pressure turbine 8. In fact, the overall efficiency of the propulsion system is primarily regulated by the propulsion efficiency, which is favorably 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 propulsion force is composed of the secondary air flow F2 of the propulsion system 1, the kinetic energy of which is primarily influenced by the compression experienced by the secondary air flow F2 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 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 ).

[0086] 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).

[0087] The fan section 2 may be of ducted or unducted type. In the case of a ducted fan section 2, the fan section 2 comprises a fan casing 12 in which the fan rotor 9 is accommodated.

[0088] The ducted fan section 2 includes a fan rotor 9 extending upstream of the fan stator. The fan stator's blades are typically referred to as outlet guide vanes (OGVs) and have a fixed setting relative to the fan stator's hub. Furthermore, the circumferential speed at the tips 21 of the fan rotor 9 blades can also be between 260 m / s and 400 m / s. The blades 14 of the fan rotor 9 can be fixed or have a variable setting.

[0089] 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 referred to as an "open rotor" or "unducted fan". The propulsion system 1 may comprise two unducted and counter-rotating fan rotors 9. Such a propulsion system 1 is referred to as a "Contra-Rotating Open Rotor (CROR)" or an "Unducted Double Fan (UDF)". One or more fan rotors 9 may be arranged at the rear of the main body 3 to form a push-type structure, or at the front of the main body 3 to form a pull-type structure. As a variant, the propulsion system 1 may comprise a single unducted fan rotor 9 and an unducted fan stator 16 (straightener). Such a propulsion system 1 is 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.

[0090] Removing the fairing surrounding the fan section 2 makes it possible to increase the bypass ratio very significantly, without the propulsion system 1 being adversely affected by the mass of the casing or nacelle intended to surround the fan section 2. The peripheral speed at the tips 21 of the blades 14 of the fan rotor(s) 9 may also be between 210 m / s and 260 m / s.

[0091] The reduction mechanism 19 may comprise, for example, a reduction mechanism 19 having a planetary gear train, a single-stage or two-stage reduction mechanism 19 of a "planetary" or "star type" type, according to the terminology encountered by those skilled in the art. According to a first variant, the reduction mechanism 19 may be of a star type ( Figure 3), 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 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 housing of the compressor sections 4, 5).

[0092] 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 .

[0093] 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, for example, approximately 3.0. In the case of a propulsion system 1 including an unducted fan rotor, the reduction ratio may be between 9.0 and 11.0.

[0094] The limit speed of the low-pressure shaft 11 is between 8500 rpm and 12000 rpm, for example 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 European Certification Regulation (EASA) CS-E 740 (or according to US Certification Regulation 14-CFR Part 33)). The limit speed corresponds to the maximum rotational speed when the propulsion system is 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).

[0095] In order to optimize the performance of the propulsion system 1 , in particular the footprint of the propulsion system 1 and the overall dynamics of the propulsion system 1 , the average outer radius of the high-pressure compressor satisfies the following formula:

[0096]

[0097] Where: R ext_5 is the average outer radius of the high-pressure compressor, in millimeters (mm);

[0098] 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 ;

[0099] BPR is the bypass ratio of propulsion system 1;

[0100] T e is the maximum inlet temperature of the low-pressure turbine 8, in degrees Celsius (°C);

[0101] N is the rotational speed of the high-pressure shaft 10 in revolutions per minute (tr / min);

[0102] S7 is the average surface area of ​​the high-pressure turbine 7, in square millimeters (mm 2 );as well as

[0103] A=1(tr / min) 2 .mm / ℃ and B=8977(tr / min) 2 *mm 2 .

[0104] The average outer radius R of the high-pressure compressor 5 ext_5is equal to the arithmetic mean of the outer radius R1 of all the rotors (blade wheels) of the high-pressure compressor 5. In a given stage, the outer radius R1 of the rotor corresponds to the distance between the tip 5e of the blade of the rotor 5a and the axis of rotation X, this distance being measured halfway between the leading edge 5c and the trailing edge 5d of the blade of the rotor 5a (at the tip 5e) (50% of the chord at the tip 5e).

[0105] The average surface S7 of the high-pressure turbine 7 is equal to the arithmetic mean of the flow path surfaces at 50% of the chord at the tip 7e of all the rotors 7b (blade wheels) of the high-pressure turbine 7. The surface of the flow path of a given rotor 7b of the high-pressure turbine 7 is equal to π×(R ext_7 2 -R int_7 2 ), where R ext_7 and R int_7 Corresponding to the average outer radius and average inner radius of the high pressure turbine 7 respectively. The average outer radius R of the high pressure turbine 7 ext_7 (or average inner radius R int_7 ) is equal to the arithmetic mean of the outer radius R3 (or inner radius R4) of the rotor 7b (blade wheel) of the high-pressure turbine 7. In a given stage, the outer radius R3 of the rotor corresponds to the distance between the tip 7e of the blade of the rotor 7b and the axis of rotation X, which is measured at the middle between the leading edge 7c and the trailing edge 7d of the blade of the rotor 7b (at the tip 7e) (50% of the chord at the tip 7e); the inner radius R4 of the rotor corresponds to: in a plane passing through 50% of the chord at the tip 7e, the distance between the axis of rotation X and the following surface, so that the outer radius R3 and the inner radius R4 are measured in the same plane, which surface delimits the flow path in the rotor 7b on the radial inside.

[0106] The average surface S7 of the high pressure turbine may be, for example, between 0.035 m 2 to 0.1m 2 between.

[0107] As mentioned above, the bypass ratio BPR, temperature T e and the rotation speed N are determined under takeoff conditions. On the other hand, the average outer radius R ext_5 and diameter D9 are determined when the propulsion system 1 is cold.

[0108] The inlet temperature of the low-pressure turbine is T e The temperature can be between 950°C and 1230°C. The inlet cross section of the high pressure turbine can be between 0.08m 2 Up to 0.35m 2 between.

[0109] When the average outer radius R ext_5When formula (1) is satisfied, the radial and longitudinal space occupied by the high-pressure compressor 5 and the mass of the high-pressure compressor 5 are reduced with similar performance. In fact, the power density of the high-pressure compressor 5 is related to the inlet flow rate of the low-pressure compressor 4 and the inlet temperature T of the low-pressure turbine 8. e It is known that the inlet flow of the low-pressure compressor 4 is a function of the bypass ratio BPR of the propulsion system 1 and the diameter D9 of the fan rotor 9 (for the same ratio D / S, where S corresponds to the inlet section of the fan rotor 9). Therefore, by taking the average outer radius R ext_5 The dimensioning of is determined as a function of the fan diameter D9 and the bypass ratio BPR, making it possible to reduce the average outer radius of the high-pressure compressor 5 while ensuring that the power density of the high-pressure compressor 5 is sufficient to obtain an efficient propulsion system 1 (and therefore good energy performance).

[0110] The mean outer radius R is also determined by taking into account the mechanical load of the high-pressure turbine 7 ext_5 The size of the mechanical load is proportional to the square of the speed of the high-pressure shaft 10 (N 2 ) is proportional to the product of the average surface S7 of the high-pressure turbine 7.

[0111] Therefore, the average outer radius R is determined according to formula (1): ext_5 The dimensions of the high-pressure compressor 5 make it possible to reduce the footprint and mass of the high-pressure body without mechanically compromising the high-pressure turbine 7 or affecting the efficiency of the propulsion system 1. The reduction in the radial footprint of the high-pressure compressor 5 also makes it possible to facilitate the integration of the propulsion system 1 into the aircraft 100 and reduce the drag of the propulsion system 1.

[0112] In particular, the average outer radius (R ) of the high-pressure compressor 5 can be obtained by increasing the bypass ratio BPR of the propulsion system 1 and by reducing the pressure ratio of the fan rotor 9 and, therefore, by increasing the efficiency of the propulsion system 1 . ext_5 ) for the ducted fan section 2, the bypass ratio BPR of the propulsion system 1 may be greater than or equal to 10, for example, between 10 and 35 (inclusive), for example, between 10 and 18 (inclusive). The fan pressure ratio may also be between 1.20 and 1.45. For the unducted fan section 2, the bypass ratio BPR of the propulsion system 1 may be greater than or equal to 40, for example, between 40 and 80 (inclusive). The fan pressure ratio may also be between 1.05 and 1.20.

[0113] Alternatively or additionally, this reduction can be obtained by improving the aerodynamic performance of the high-pressure body, by increasing the allowed Mach for the same efficiency for the same aerodynamic load and / or by increasing the aerodynamic load for the same efficiency (possibly compensated by reducing the Mach in the path of the high-pressure compressor 5), in order to increase the power of the high-pressure turbine 7 and the mechanical resistance of the high-pressure turbine 7. For example, a reduction in leakage in the high-pressure turbine 7 (by maintaining the clearances) makes it possible to increase the efficiency of the compression carried out in the high-pressure compressor 5 and of the expansion carried out in the high-pressure turbine 7.

[0114] When the average outer radius (R ext_5 ) When formula (1) is satisfied, the high-pressure turbine 7 can also rotate at a sufficiently high speed (in revolutions per minute) (and at an acceptable mechanical load) so that the high-pressure compressor 5 can achieve a high power density and a compression ratio greater than 21 (at takeoff rating) without increasing the number of stages of the high-pressure compressor 5 (and therefore by limiting the longitudinal footprint of the high-pressure compressor 5). The speed of the high-pressure shaft 10 can be between 15,000 rpm and 27,000 rpm. The high-pressure turbine 7 is a two-stage turbine, the high-pressure compressor 5 is axial, and the high-pressure compressor 5 can include at least eight stages and at most eleven stages, for example nine stages.

[0115] An overall compression ratio of the propulsion system 1 is also obtained, 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 can be greater than or equal to 40 and less than or equal to 70, for example greater than or equal to 44 and less than or equal to 55.

[0116] For example, the average outer radius R ext 5 also satisfies the following formula:

[0117]

[0118] Where: C = 7668 (tr / min) 2 *mm 2 .

[0119] When the average outer radius R ext_5 When equation (2) is also satisfied, the high-pressure compressor 5 remains large enough to compress the air at the inlet of the combustion chamber 6, thereby contributing to the overall compression ratio of the propulsion system 1. In addition, the aerodynamic loads and Mach in the flow path remain compatible with the efficiency and operability targets of the high-pressure compressor 5.

[0120] For the average outer radius R ext_5For a high-pressure compressor 5 that satisfies formula (1) (and formula (2) where appropriate), the diameter D9 of the fan rotor can be between 2032 mm (80 inches) and 4699 mm (185 inches) (including 2032 mm and 4699 mm). When the fan rotor 9 is of ducted type, the diameter D9 is, for example, between 2159 mm (85 inches) and 3048 mm (120 inches) (including 2159 mm and 3048 mm), for example, approximately 2286 mm (90 inches), which enables the propulsion system 1 to be integrated in a conventional manner, in particular, under the wing of an aircraft. When the fan rotor 9 is of non-ducted type, the diameter D9 is, for example, greater than or equal to 2540 mm (100 inches), for example, between 3048 cm (120 inches) and 3962 mm (156 inches).

[0121] Where appropriate, in order to achieve a compromise between the radial compactness of the high-pressure compressor 5 and the aerodynamic loads of the high-pressure compressor 5 , the average outer radius R ext_5 The following formula is also satisfied:

[0122] (E*n5+F)*R ext_7 <R ext_5 <(E*n5+G)*R ext_7 (3)

[0123] Wherein: n5 is the number of stages of the high-pressure compressor 5; and

[0124] E=-23.9*10 -3 Square millimeters (mm 2 ), F = 0.96 and G = 1.06, F and G are dimensionless.

[0125] For the same number of stages, compliance with formula (3) enables a radially more compact high-pressure compressor 5 to be obtained without compromising the efficiency of the high-pressure compressor 5. In fact, the average outer radius R of the high-pressure compressor 5 is ext_5 The smaller the average outer radius R is, the more compact the high pressure compressor 5 is. ext_5 The reduction of has the effect of increasing the aerodynamic load of the high-pressure compressor 5 (and thus reducing the efficiency of the high-pressure compressor 5). However, formula (1) makes it possible to reach an acceptable compromise between radial compactness and efficiency of the high-pressure compressor 5 and increase the mass and specific consumption of the propulsion system 1. In addition, taking into account the average outer radius R of the high-pressure turbine 7 ext_7 In formula (1), the mechanical load of the high-pressure turbine 7 (which is proportional to the product of the square of the rotational speed of the high-pressure shaft 10 and the average surface S7 of the high-pressure turbine 7) is still acceptable.

[0126] A high-pressure compressor 5 satisfying formula (1) and, where appropriate, formula (2) and formula (3) may have a hub-to-tip ratio at the inlet of the high-pressure compressor 5, which corresponds to the ratio between the outer radius R1 of the high-pressure compressor 5 and the inner radius R2 of the rotor 5b of the high-pressure compressor 5, and is between 0.41 and 0.60.

[0127] Here, the outer radius R1 and the inner radius R2 are measured in a plane perpendicular to the axis of rotation X, which plane intersects the leading edge and blade roots of the most upstream rotor 5b of the high-pressure compressor 5 (i.e., the first stage of the high-pressure compressor 5). The inner radius R2 of the high-pressure compressor 5 corresponds to the distance between the outer radial surface of the hub of the rotor 5b (which radially innermost defines the flow path in the rotor 5b) and the axis of rotation X in this plane.

[0128] When the high-pressure compressor 5 includes a rotating wheel (wholly or partially a centrifugal compressor), the rotating wheel is counted as two stages when calculating the number n5 of stages of the high-pressure compressor 5 .

[0129] This high-pressure compressor 5 has an optimized outlet cross-section. In fact, the smaller the hub-to-tip ratio, the smaller the outer diameter of the high-pressure compressor 5 (for the same cross-section). Thus, a hub-to-tip ratio between 0.77 and 0.90 is consistent with the optimized average outer radius R as described above. ext_5 In combination with the rotational speed, this makes it possible to obtain a more efficient high-pressure compressor 5 in a suitable occupied space (due to the adapted rotational speed of the high-pressure turbine 7 ), while optimizing the surface for compressing the gas at the inlet of the combustion chamber 6 .

[0130] having an average outer radius R in the interval defined above ext_5 The twin-body propulsion system 1 may particularly include a high-pressure turbine 8 including at least three stages and at most five stages, and a low-pressure compressor 4 including at least two stages and at most four stages.

[0131] Comparative Example 1: Ducted Propulsion System

[0132] The engine 1 is a twin-body propulsion system comprising a ducted fan section 2 corresponding to the current technical standards (at the filing date of the present application) for which improvement is sought.

[0133] The engine 2 is a twin-body propulsion system 1, which includes a ducted fan section 2 in accordance with the teachings of the present application, the ducted fan section 2 having an average radius of the high-pressure compressor between the maximum value and the minimum value defined by the relationships (1), (2) and (3) as defined above.

[0134]

[0135]

[0136] The high-pressure compressor of the engine 1 has a large average outer radius that does not satisfy the above-defined formulas (1) and (3).

[0137] In contrast, the high-pressure compressor 5 of engine 2 has a moderate average outer radius that satisfies equations (1) to (3). Therefore, the high-pressure body of engine 2 is more compact than that of engine 1 without causing mechanical loads on the high-pressure turbine 7, while ensuring that the aerodynamic load of the high-pressure compressor 5 remains moderate to achieve an efficient propulsion system 1 (and therefore provide good energy performance).

[0138] To convert from (reference) engine 1 to engine 2 (in accordance with the present disclosure), the fan diameter D and the bypass ratio BPR are increased, which makes it possible to increase propulsion efficiency while maintaining comparable fan thrust, given a reduced pressure ratio of the fan section 2. In addition, 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.

[0139] Comparative Example 2: Unducted Propulsion System

[0140] The engine 3 is a twin-body propulsion system comprising an unducted fan section 2 corresponding to the current technical standard (at the filing date of the present application) for which improvement is sought.

[0141] The engine 4 is a twin-body propulsion system 1, which includes a non-ducted fan section 2 in accordance with the teachings of the present application, and the non-ducted fan section 2 has an average radius of the high-pressure compressor between the maximum value and the minimum value defined by the relationships (1), (2) and (3) as defined above.

[0142]

[0143]

[0144] The high-pressure compressor of the engine 3 has a large average outer radius that does not satisfy the formula (1) and the formula (3) defined above.

[0145] In contrast, the high-pressure compressor 5 of the engine 4 has a moderate average outer radius that satisfies equations (1) to (3). As a result, the high-pressure body of the engine 4 is more compact than that of the engine 3, and the mechanical load of the high-pressure turbine 7 is lower, while ensuring that the power density of the high-pressure compressor 5 remains sufficient to obtain an efficient propulsion system 1 (and therefore provide good energy performance).

[0146] To convert from (reference) engine 3 to engine 4 (in accordance with the present disclosure), the fan diameter D9 and the bypass ratio BPR are reduced, which enables improved integration of engine 2. However, the pressure ratio of the fan section of engine 2 is slightly increased (while remaining below 1.45) to maintain equivalent thrust. In addition, the overall compression ratio is increased without increasing the inlet temperature of the high-pressure turbine 7, which improves the efficiency of the body without increasing the thermal load of the low-pressure turbine 8. Finally, as long as the temperature of the low-pressure turbine 8 remains stable, the mechanical load of the low-pressure turbine 8 can be increased to reduce the number of stages of the low-pressure turbine 8.

[0147] The increase in the power density of the high-pressure compressors of Engine 2 and Engine 4 can be achieved as follows:

[0148] -Increase the compression ratio of the high-pressure compressor;

[0149] - reducing the diameter of the low-pressure shaft while maintaining acceptable supercritical dynamic conditions (first deformation mode) within the operating range. In the operating area where the dynamic conditions of the low-pressure shaft 11 can generally be controlled, for a propulsion system including a reduction gear 19, the first deformation mode of the low-pressure shaft 11 occurs at a transient speed corresponding, for example, to a speed lower than the idle speed and is therefore unlikely to damage the propulsion system 1. Therefore, by modifying the internal geometry of the low-pressure shaft and increasing the distance between the bearings downstream of the low-pressure turbine, a transient speed (second deformation mode outside the operating range) is selected that occurs only for a very short time during the operation of the propulsion system 1;

[0150] - Better integration of the upstream bearing of the high-pressure body combined with a reduction in the diameter of the low-pressure shaft allows for a reduction in the internal path. This better integration can be achieved, for example, by using rolling elements made of ceramic material, which increases the contact pressure and reduces heat rejection.

[0151] Alternatively, instead of reducing the diameter of the low-pressure shaft, the shape of the low-pressure shaft can be modified to give it a bottle-like shape: specifically, the radius of the low-pressure shaft below the high-pressure turbine is changed to maintain its dynamic conditions, and the radius of the low-pressure shaft upstream of the combustion chamber is reduced to facilitate integration of the internal path of the high-pressure compressor. Reducing the aerodynamic load on the high-pressure turbine also increases the diameter of the high-pressure turbine and facilitates the design of the high-pressure turbine disk.

Claims

1. An aviation propulsion system (1), comprising: - a fan rotor (9), said fan rotor being connected to a fan shaft (20); a first turbine (8) configured to drive the fan rotor (9) 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 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 outer radius of the second compressor (5) satisfies the following formula: Where: R ext_5 is the average outer radius of the second compressor (5), in millimeters (mm); D9 is the diameter of the fan rotor (9) in millimeters (mm), and D9 is measured in a plane perpendicular to the axis of rotation (X) of the fan rotor (9) 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 first turbine (8) when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C); N is the rotational speed of the second shaft (10) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, in revolutions per minute (tr / min); S7 is the average surface area of ​​the second turbine (7), expressed in square millimeters (mm 2 ) means; and A = 1 (tr / min) 2 .mm / ℃ and B=8977(tr / min) 2 *mm 2 .

2. The propulsion system (1) according to claim 1, wherein The average outer radius of the second compressor (5) also satisfies the following formula: Where: C = 7668 (tr / min) 2 *mm 2 .

3. The propulsion system (1) according to claim 1 or 2, wherein: The average outer radius of the second compressor (5) also satisfies the following formula: R ext_5 >(E*n5+F)*R ext_7 Wherein: n5 is the number of stages of the second compressor (5); R ext_7 is the average outer radius of the second turbine (7), in millimeters (mm); and E=-23.9*10 -3 Square millimeters (mm 2 ) and F = 0.

96.

4. The propulsion system (1) according to any one of claims 1 to 3, wherein: The average outer radius of the second compressor (5) also satisfies the following formula: R ext_5 <(E*n5+G)R ext_7 Wherein: n5 is the number of stages of the second compressor (5); R ext_7 is the average outer radius of the second turbine (7), in millimeters (mm); and E=-23.9*10 -3 Square millimeters (mm 2 ) and G=1.

06.

5. The propulsion system (1) according to any one of claims 1 to 4, wherein: The second turbine (7) is a two-stage turbine, and the second compressor (5) is axial.

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

7. The propulsion system (1) according to any one of claims 1 to 6, wherein: The diameter of the fan rotor (9) is between 2032 mm (80 inches) and 4699 mm (185 inches) inclusive, preferably between 2159 mm (85 inches) and 3048 mm (120 inches) inclusive, for example, about 2286 mm (90 inches).

8. The propulsion system (1) according to any one of claims 1 to 7, wherein: The fan section (2) is of a ducted type, and the bypass ratio of the propulsion system (1) is greater than or equal to 10, such as between 10 and 35 and including 10 and 35, such as between 10 and 18 and including 10 and 18.

9. The propulsion system (1) according to any one of claims 1 to 7, wherein: The fan section (2) 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, inclusive.

10. The propulsion system (1) according to any one of claims 1 to 9, wherein: The hub-to-tip ratio at the inlet of the second compressor (5) is between 0.41 and 0.

60.

11. The propulsion system (1) according to any one of claims 1 to 10, wherein: The rotation speed of the second shaft (10) is greater than or equal to 15,000 revolutions per minute (tr / min) and less than or equal to 27,000 revolutions per minute (tr / min).

12. The propulsion system (1) according to any one of claims 1 to 11, 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.

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 turbine also drives a first compressor (4) via the first shaft (11), the first compressor (4) comprising at least two stages and at most four stages.

15. The propulsion system (1) according to any one of claims 1 to 14, wherein: The fan section (2) also has a fan pressure ratio, which corresponds to the pressure ratio between the outlet of the fan rotor (9) and the inlet of the fan rotor (9), which is less than or equal to 1.45, for example less than or equal to 1.

30.

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

17. A method for determining the size of a propulsion system (1), the propulsion system comprising a speed 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), wherein: The average outer radius of the second compressor (5) satisfies the following formula: Where: R ext_5 is the average outer radius of the second compressor (5), in millimeters (mm); D9 is the diameter of the fan rotor (9) in millimeters (mm), and D9 is measured in a plane perpendicular to the axis of rotation of the fan rotor (9) 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 first turbine (8) when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, expressed in degrees Celsius (°C); N is the rotational speed of the second shaft (10) and is measured when the propulsion system (1) is at rest, takeoff rating, standard atmospheric pressure and sea level conditions, in revolutions per minute (tr / min); S7 is the average surface area of ​​the second turbine (7), expressed in square millimeters (mm 2 ) means; and A = 1 (tr / min) 2 .mm / ℃ and B=8977(tr / min) 2 *mm 2 .

18. The size determination method according to claim 17, wherein: The average outer radius of the second compressor (5) satisfies the following formula: Where: C = 7668 (tr / min) 2 *mm 2 .

19. The size determination method according to claim 17 or 18, wherein: The average outer radius of the second compressor (5) also satisfies the following formula: R ext_5 >(E*n5+F)*R ext_7 Wherein: n5 is the number of stages of the second compressor (5); R ext_7 is the average outer radius of the second turbine (7), in millimeters (mm); and E=-23.9*10 -3 Square millimeters (mm 2 ) and F = 0.

96.

20. The size determination method according to any one of claims 17 to 19, wherein: The average outer radius of the second compressor (5) also satisfies the following formula: R ext_5 <(E*n5+G)*R ext_7 Wherein: n5 is the number of stages of the second compressor (5); R ext_7 is the average outer radius of the second turbine (7), in millimeters (mm); and E=-23.9*10 -3 Square millimeters (mm 2 ) and G=1.

06.

21. A method for manufacturing an aviation propulsion system, comprising the steps of: - sizing the aerospace propulsion system according to the method of any one of claims 17 to 20; and - Manufacturing of said aviation propulsion system.

Images