Aerial propulsion system with improved propulsion efficiency

Abstract

An aero propulsion system (1) comprising an open fan rotor (9) having an aerodynamic load parameter kaero greater than or equal to 0.07 and less than or equal to 0.12, and a mechanical load parameter kmeca greater than or equal to 11.5 * 106 tr.m2. Min-1 and less than or equal to 20.0 * 106 tr.m2. Min-1, where the aerodynamic load parameter kero is greater than or equal to 0.07 and less than or equal to 0.12, and the mechanical load parameter kmeca is greater than or equal to 11.5 * 106 tr.m2. Min-1 and less than or equal to 20.0 * 106 tr.m2. Min-1. The aerodynamic load kaero and the mechanical load kmeca are defined by the following formulas: formula (1) # imgabs0 #, wherein FN is the thrust of the propulsion system (1); bPR is a bypass ratio; f is equal to 39.7; re and Ri are the outer radius and the inner radius of the fan rotor; omega is the limit speed of the fan rotor; and S is the surface area of the flow stream at the inlet of the fan rotor (9).

Description

Technical Field

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

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

[0003] Technological research has significantly improved the environmental performance of aircraft. Applicants consider influencing factors at all stages of design and development to obtain aviation components and products that consume less energy, are more environmentally friendly, and have a moderate impact on the environment when integrated and used in civil aviation, with the goal of improving the energy efficiency of aircraft.

[0004] In order to improve the propulsion efficiency of the propulsion system and reduce the specific consumption of the propulsion system and the noise emitted by the fan section, a propulsion system with a high bypass ratio (BPR, the bypass ratio corresponds to the ratio between the secondary air flow and the main air flow) has been proposed. In order to achieve such a bypass ratio, the fan section can be separated from the low-pressure turbine so that the corresponding rotational speeds of the fan section and the low-pressure turbine can be optimized independently. Usually, the separation is achieved by a reduction mechanism arranged between the upstream end of the low-pressure shaft and the rotor of the fan section. The rotor of the fan section is then driven by the low-pressure shaft via the reduction mechanism at a speed lower than the speed of the low-pressure shaft.

[0005] This separation enables the rotational speed of the fan rotor to be reduced. Furthermore, in order to reduce the pressure ratio of the fan rotor and improve the bypass ratio of the propulsion system, it has been proposed to increase the diameter of the fan rotor (and thus reduce the circumferential speed of the fan blades). This results in a reduction in the centrifugal force in the fan rotor. However, increasing the fan diameter has a negative impact on the mass and drag of the fan section, and thus on the propulsion system, and increases the torque applied to the fan rotor by the fan shaft. Summary of the Invention

[0006] One object of the present application is to optimize the performance of a propulsion system while maintaining acceptable mechanical and aerodynamic loads on the fan section of the propulsion system.

[0007] To this end, a first aspect of the present invention provides an aviation propulsion system, comprising:

[0008] - a drive shaft, the drive shaft being rotatable about an axis of rotation;

[0009] -Fan shaft;

[0010] a fan section, the fan section including an unducted fan rotor rotated by a fan shaft, the fan rotor including a plurality of blades;

[0011] - a speed reduction mechanism coupling the drive shaft and the fan shaft so as to drive the fan shaft at a rotational speed lower than that of the drive shaft;

[0012] The propulsion system is configured such that the fan rotor has an aerodynamic load parameter greater than or equal to 0.07 and less than or equal to 0.12, and a load factor greater than or equal to 11.5×10 6 tr.m 2 .min -1 and less than or equal to 20.0×10 6 tr.m 2 .min -1 Mechanical load parameters,

[0013] The aerodynamic load and mechanical load are defined by the following formulas:

[0014]

[0015] Where: k aero is the aerodynamic load parameter;

[0016] k meca is the mechanical load parameter;

[0017] FN is the thrust produced by the propulsion system when the propulsion system is at rest, takeoff speed, standard atmospheric pressure, and sea level conditions, expressed in Newtons;

[0018] BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is at rest, takeoff speed, standard atmospheric pressure and sea level conditions;

[0019] n is the number of blades in the fan rotor;

[0020] F=39.7N -1.35 .(tr / min) -2 .m -5 ;

[0021] R e is the outer radius of the fan rotor and corresponds to the distance between the axis of rotation and the intersection between the tips and leading edges of the blades of the fan rotor, expressed in meters (m);

[0022] R iis the inner radius of the fan rotor and corresponds to the distance between the axis of rotation and the intersection of the leading edges of the blades of the fan rotor and the surface of the fan rotor which delimits the flow stream in the fan rotor radially inside, expressed in meters (m); and

[0023] ω is the fan rotor's limiting speed, expressed in revolutions per minute (tr / min); and

[0024] S is the surface area of ​​the flow at the inlet of the fan rotor, which is equal to π×(R e 2 -R i 2 ), in square meters (m 2 )express.

[0025] Some preferred but non-limiting features of the propulsion system according to the first aspect are the following features taken alone or in combination:

[0026] The fan rotor has an outer radius between 40 inches (101.6 cm) and 92.5 inches (233.7 cm), inclusive, such as between 60 inches (152.4 cm) and 85 inches (215.9 cm), inclusive, such as approximately 478 inches (198.1 cm);

[0027] -The fan rotor has a hub-to-tip ratio between 0.22 and 0.32;

[0028] - the fan rotor comprises at least twelve fan blades and at most eighteen fan blades, for example at least fourteen fan blades and at most sixteen fan blades;

[0029] - The reduction ratio of the reduction mechanism is greater than or equal to 5.0 and less than or equal to 11.0;

[0030] - a propulsion system bypass ratio greater than or equal to 40, such as between 40 and 80 inclusive;

[0031] - the peripheral speed at the tips of the blades of the fan rotor is between 210 m / s and 260 m / s when the propulsion system is at rest, take-off speed, standard atmospheric pressure and sea level conditions;

[0032] The fan section further comprises a fan compression ratio, the fan compression ratio corresponding to a pressure ratio between an outlet of the fan rotor and an inlet of the fan rotor, the pressure ratio being less than or equal to 1.45, for example less than or equal to 1.30;

[0033] - the propulsion system is sized so that the thrust of the propulsion system is between 18,000 lbf and 51,000 lbf when the propulsion system is at rest, takeoff speed, and standard atmospheric pressure conditions;

[0034] - the propulsion system further comprises a drive turbine and a compressor directly connected via a drive shaft, wherein the drive turbine comprises at least three stages and at most five stages;

[0035] - The compressor comprises at least two stages and at most four stages;

[0036] - the propulsion system further comprises a high-pressure turbine and a high-pressure compressor connected via a high-pressure shaft, the high-pressure shaft rotating faster than the drive shaft, the high-pressure turbine being two-stage; and / or

[0037] The high-pressure compressor comprises at least eight stages and at most eleven stages.

[0038] According to a second aspect, an aircraft is proposed, comprising at least one propulsion system according to the first aspect, the propulsion system being attached to the aircraft via a mast.

[0039] According to a third aspect, a method for sizing or manufacturing a propulsion system is provided, the propulsion system including a reduction mechanism coupling a drive shaft and an unducted fan rotor such that the fan rotor is driven at a speed less than the speed of the drive shaft; wherein the fan rotor has an aerodynamic load parameter greater than or equal to 0.07 and less than or equal to 0.12, and a speed greater than or equal to 11.5×10 6 tr.m 2 .min -1 and less than or equal to 20.0×10 6 tr.m 2 .min -1 Mechanical load parameters,

[0040] The aerodynamic load and mechanical load are defined by the following formulas:

[0041]

[0042] Where: k aero is the aerodynamic load parameter;

[0043] k meca is the mechanical load parameter;

[0044] FN is the thrust produced by the propulsion system when the propulsion system is at rest, takeoff speed, standard atmospheric pressure, and sea level conditions, expressed in Newtons (N);

[0045] BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is at rest, takeoff speed, standard atmospheric pressure and sea level conditions;

[0046] n is the number of blades in the fan rotor;

[0047] F=39.7N -1.35 .(tr / min) -2 .m -5 ;

[0048] R e is the outer radius of the fan rotor and corresponds to the distance between the axis of rotation and the intersection between the tips and leading edges of the blades of the fan rotor, expressed in meters (m);

[0049] R i is the inner radius of the fan rotor and corresponds to the distance between the axis of rotation and the intersection of the leading edges of the blades of the fan rotor and the surface of the fan rotor which delimits the flow stream in the fan rotor radially inside, expressed in meters (m); and

[0050] ω is the fan rotor's limiting speed, expressed in revolutions per minute (tr / min); and

[0051] S is the surface area of ​​the flow at the inlet of the fan rotor, which is equal to π×(R e 2 -R i 2 ), in square meters (m 2 )express.

[0052] According to a fourth aspect, a method for manufacturing an aviation propulsion system is provided, comprising the following steps:

[0053] - sizing an aerospace propulsion system according to the sizing method of the third aspect; and

[0054] -Manufacture of aviation propulsion systems. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] 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 with reference to the accompanying drawings, in which:

[0056] 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 unducted; Figure 2 is a schematic cross-sectional view of an example of a speed reduction mechanism according to a first alternative embodiment;

[0057] Figure 3is a schematic cross-sectional view of an example of a speed reduction mechanism according to a second alternative embodiment;

[0058] Figure 4 is an example of an aircraft that may include at least one propulsion system according to the first or second embodiment of the invention;

[0059] Figure 5 is a flow chart illustrating an example of steps in a sizing or manufacturing method according to an embodiment of the present invention.

[0060] In all figures, similar elements have the same reference numerals. DETAILED DESCRIPTION

[0061] The propulsion system 1 has a main direction extending along a longitudinal axis X. When the propulsion system 1 is in operation, the propulsion system 1 comprises, from upstream to downstream in the direction of gas flow in the propulsion system 1, a fan section 2 and a main body 3 (commonly referred to as a "gas generator"), which includes compressor sections 4 and 5, a combustion chamber 6, and turbine sections 7 and 8. Here, the propulsion system 1 is an aircraft propulsion system 1, which is configured to be fixed to an aircraft 100 via a pylon (or mast).

[0062] The compressor sections 4 and 5 include a series of stages, each of which includes a moving impeller (rotor) 4a and 5a rotating in front of a stationary impeller (stator) 4b and 5b. The turbine sections 7 and 8 also include a series of stages, each of which includes a stationary impeller (stator) 7b and 8b, with the moving impeller (rotor) 7a and 8a rotating behind the stationary impeller (stator).

[0063] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, corresponding to the rotation of the shaft of the gas generator, and the radial direction is a direction perpendicular to and passing through this axis X. On the other hand, the circumferential (or transverse or tangential) direction corresponds to a direction perpendicular to the longitudinal axis X but not passing through the longitudinal axis X. Unless otherwise specified, the terms inner (inner) and outer (outer) 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.

[0064] In operation, the air flow F entering the propulsion system 1 is divided into a primary air flow F1 and a secondary air flow F2 , which flow from upstream to downstream in the propulsion system 1 .

[0065] 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.

[0066] The main air flow F1 flows in the main duct within the main body 3, thereby passing through the compressor sections 4 and 5, the combustion chamber 6, and the turbine sections 7 and 8 in sequence. In the combustion chamber 6, the main air flow F1 is mixed with fuel to be used as an oxidizer. The main air flow F1 flows through the turbine sections 7 and 8, which receive energy from the combustion chamber 6. This causes the rotors of the turbine sections 7 and 8 to rotate. The rotation of the rotors of the turbine sections 7 and 8 in turn rotates the rotors of the compressor sections 4 and 5 and the rotor portion 9 of the fan section 2.

[0067] In a dual-body propulsion system 1, the compressor sections 4 and 5 may include a low-pressure compressor 4 and a high-pressure compressor 5. The turbine sections 7 and 8 may include a high-pressure turbine 7 and a low-pressure turbine 8. The high-pressure turbine 7 rotates the rotor of the high-pressure compressor 5 via a high-pressure shaft 10. The low-pressure turbine rotor 8 rotates the rotor of the low-pressure compressor 4 and the rotor portion 9 of the fan section 2 via a low-pressure shaft 11. Thus, the main body 3 comprises 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 higher than that of the low-pressure main body. In a three-shaft propulsion system 1, the turbine sections 7 and 8 also include an intermediate turbine, located between the high-pressure turbine 7 and the low-pressure turbine 8 and configured to drive the rotor of the low-pressure compressor 4 via the intermediate shaft. The fan rotor 9 and the rotor of the high-pressure compressor 5 are still driven by the low-pressure shaft 11 and the high-pressure shaft 10, respectively.

[0068] The low-pressure shaft 11 is typically housed within the high-pressure shaft 10 over a portion of its length and is coaxial with the high-pressure shaft 10. The low-pressure shaft 11 and the high-pressure shaft 10 can rotate in the same direction, i.e., be driven in the same direction about the longitudinal axis X. Alternatively, the low-pressure shaft 11 and the high-pressure shaft can rotate in opposite directions, i.e., be driven in opposite directions about the longitudinal axis X. Where 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.

[0069] The fan section 2 includes at least one fan rotor 9, which is rotatable relative to the stator portion of the propulsion system via the turbine sections 7 and 8. Each fan rotor 9 includes a hub 13 and blades 14 extending radially from the hub 13. The blades 14 of each rotor 9 have a variable pitch. The blade root 14 of each rotor 9 is pivotally mounted along a pitch axis and connected to a pitch changing mechanism 15 installed in the propulsion system 1. The pitch is adjusted by the pitch changing mechanism 15 according to the flight phase.

[0070] The fan section 2 may further comprise a fan stator 16 or a flow straightener comprising blades 17 mounted on a hub 18 of the fan stator 16, the blades 17 having the function of straightening the secondary air flow F2 flowing out of the fan rotor 9. This is particularly true when the fan section 2 is unducted (see Figure 2 ), the blades 17 of the fan stator 18 can be fixed or have a variable pitch relative to the hub 18. In a manner similar to the rotor blades 14, the roots of the variable-pitch stator blades 17 are pivotally mounted along the pitch axis X and are connected to a pitch changing mechanism 15a, usually separate from the pitch changing mechanism of the fan rotor 9, by which the pitch is adjusted according to the flight phase.

[0071] To improve the propulsion efficiency of the propulsion system 1 and reduce the specific consumption of the propulsion system 1 and the noise emitted by the fan section 2, the propulsion system 1 has a high bypass ratio. The term "high bypass ratio" shall refer to a bypass ratio greater than or equal to 10, for example, between 40 and 80 (inclusive). To calculate the bypass ratio, the mass flow rate of the secondary air flow F2 and the mass flow rate of the primary air flow F1 are measured when the propulsion system 1 is at rest, not installed, at takeoff speed, standard atmospheric pressure (as defined by the 3rd edition of the International Civil Aviation Organization (OACI) Manual, Document 7488 / 3), and sea level conditions (referred to as sea level standard (SLS) conditions). It should be noted that in this application, parameters (pressure, flow, thrust, speed, etc.) are systematically determined under these conditions. The term "not installed" is used herein to refer to measurements taken when the propulsion system 1 is on a test bench (and not installed on the aircraft 100), in which case the measurements are more easily performed. On the other hand, when the propulsion system 1 is cooled, that is, when the propulsion system 1 has been stationary for a period of time long enough so that the components of the propulsion system 1 are at ambient temperature, distances (lengths, radii, diameters, etc.) are measured at ambient temperature (approximately 20°C), it being understood that these dimensions vary very little compared to the condition when the propulsion system 1 is at take-off speed.

[0072] The fan rotor 9 is separated from the low-pressure shaft 11 by a reduction gear 19, which is arranged between the upstream end of the low-pressure shaft 11 and the fan rotor 9, so 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 rotational speed lower than that of the low-pressure turbine 8.

[0073] This separation enables a reduction in the rotational speed and pressure ratio of the fan rotor 9 and an increase in the power extracted by the low-pressure turbine 8. The overall efficiency of the propulsion system is determined to a first order by the propulsion efficiency, which is advantageously influenced by minimizing the change in kinetic energy of the air as it passes through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, the majority of the flow generating the propulsive 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 as it passes 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, for example, less than or equal to 1.50, for example, between 1.05 and 1.45. In this case, the mean pressure is measured at the height of the blade 14 (from the surface radially delimiting the internal flow stream of the fan rotor 9 to the tip 21 of the fan blade 14 ).

[0074] The fan section 2 may be unducted.

[0075] The fan section 2 (also known as a propeller) is not surrounded by a fan casing. Since the fan section 2 is unducted, the blades 14 of the fan rotor 9 have a variable pitch. A propulsion system that includes at least one unducted fan rotor 9 is called 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 called a "Contra-Rotating Open Rotor (CROR)" or an "Unducted Double Fan (UDF)". One or more fan rotors 9 may be positioned 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. Alternatively, the propulsion system 1 may include a single unducted fan rotor 9 and an unducted fan stator 16 (flow straightener). Such a propulsion system 1 is called 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 flow straightener 16 also have a variable pitch.

[0076] Eliminating the fairing surrounding the fan section 2 enables a very significant increase in the bypass ratio, without the propulsion system 1 being adversely affected by the mass of the casing or nacelle designed to surround the fan section 2. Consequently, the bypass ratio of the propulsion system 1 including the unducted fan section 2 is greater than or equal to 40, for example, between 40 and 80 inclusive. The circumferential speed at the tips 21 of the blades 14 of the one or more fan rotors 9 may also be between 210 m / s and 260 m / s. For example, the pressure ratio of the fan may be between 1.05 and 1.20.

[0077] For example, the reduction mechanism 19 may include a reduction mechanism 19 having an epicycloid gear train, for example a single-stage or two-stage reduction mechanism of the epicycloid or planetary type, according to the terminology sometimes encountered by those skilled in the art. According to a first variant, the reduction mechanism 19 may be a planetary ("star") type ( Figure 2 ), and comprises a sun gear 19a (inlet of the reduction mechanism 19), a ring gear 19b (outlet of the reduction mechanism 19) and a series of planetary parts 19c; the sun gear 19a is centered on the rotation axis X of the reduction mechanism 19 (approximately coinciding with the longitudinal axis X) and is configured to be rotated by the low-pressure shaft 11; the ring gear 19b is coaxial with the sun gear 19a and is configured to drive the fan shaft 20 to rotate around the rotation axis X; a series of planetary parts 19c are circumferentially distributed between the sun gear 19a and the ring gear 19b around the rotation axis X, and each planetary part 19c meshes with the sun gear 19a internally and meshes with the ring gear 19b externally. This series of planetary parts 19c is mounted on a planet carrier 19d, which is fixed relative to the stator part 19e of the propulsion system 1, for example relative to the housing of the compressor sections 4, 5. According to a second variant, the reduction mechanism 19 can be a planetary type ( Figure 3 ), in which case the ring gear 19b is fixedly mounted on the stator part 19e of the propulsion system 1 and the fan shaft 20 is rotated by the planet carrier 19d (the planet carrier 19d can therefore rotate relative to the stator part 19e of the propulsion system 1, for example, relative to the housing of the compressor sections 4, 5).

[0078] Regardless of the configuration of the speed reduction mechanism 19 , the diameters of the ring gear 19 b and the planetary carrier 19 d are larger than the diameter of the sun gear 19 a , so that the rotation speed of the fan rotor 9 is lower than the rotation speed of the low-pressure shaft 11 .

[0079] 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 the European certification regulation EASACS-E 740 (or according to the US certification regulation 14-CFR Part 33.87)). The limit speed corresponds to the maximum rotational speed when the propulsion system is healthy (and possibly at the end of its life). Therefore, the low-pressure shaft 11 is likely to reach the limit speed under flight conditions. This limit speed is part of the data stated in the 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 certain certification tests (such as blade loss or rotor integrity tests).

[0080] The propulsion system 1 also has an overall compression ratio, which corresponds to the pressure ratio between the outlet pressure of the high-pressure compressor 5 and the inlet pressure of the fan rotor 9 (measured at the root of the fan rotor 9), and the overall compression ratio is greater than or equal to 40 and less than or equal to 70, for example greater than or equal to 44 and less than or equal to 55.

[0081] In order to optimize the performance of the propulsion system 1 in terms of specific fuel consumption, mass and drag, while maintaining acceptable mechanical and aerodynamic loads of the fan section 2, the propulsion system 1 is configured so that the fan rotor 9 has an aerodynamic load parameter k greater than or equal to 0.07 and less than or equal to 0.12 aero , and greater than or equal to 11.5×10 6 tr 2 .m 2 .min -2 and less than or equal to 20.0×10 6 tr 2 .m 2 .min -2 Mechanical load parameter k meca .

[0082] Aerodynamic load parameter k aero represents the aerodynamic torque applied to the fan blade 14 of the fan rotor 9 between the hub 13 and the tip 21 of the fan blade 14 and can be expressed by the following formula:

[0083]

[0084] Where: FN is the thrust generated by the propulsion system, expressed in Newtons (N);

[0085] BPR is the bypass ratio of propulsion system 1;

[0086] n is the number of blades 14 in the fan rotor 9;

[0087] ω is the limit speed (full speed) of the fan rotor 9, expressed in revolutions per minute (tr / min), which corresponds to the absolute maximum speed that the fan shaft may encounter during the entire flight;

[0088] R e is the outer radius of the fan rotor 9 and corresponds to the distance between the axis of rotation X and the intersection between the tips 21 and the leading edges 22 of the blades 14 of the fan rotor 9 , expressed in meters (m);

[0089] R i is the inner radius of the fan rotor and corresponds to the distance between the axis of rotation X and the intersection of the leading edges of the blades 14 of the fan rotor 9 and the surface of the fan rotor 9 which delimits the flow stream in the fan rotor 9 radially on the inside, expressed in meters (m); and

[0090] F=39.7N -1.35 .(tr / min) -2 .m -5 .

[0091] As previously mentioned, FN is determined when the propulsion system 1 is at rest, not installed, at takeoff speed, 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. e and R i It is determined when the propulsion system 1 is cold in order to simplify the measurement.

[0092] The leading edge 22 is configured to extend facing the gas flow entering the fan rotor 9. The leading edge 22 corresponds to a front portion of the aerodynamic profile that faces the air flow and divides the air flow into an inner arc side flow and an outer arc side flow.

[0093] Mechanical load parameter k meca Reflects the inner radius R of the fan rotor 9 i The stress applied to the fan blade at (at the base of the portion of the fan blade 14 that is configured to extend into the flow stream):

[0094] k meca =ω 2 ×S

[0095] Where S is the surface area of ​​the flow at the inlet of the fan rotor 9, which is equal to π×(R e 2 -R i 2 ), expressed in square meters.

[0096] The centrifugal force applied to the fan blades 14 of the fan rotor 9 is moderate, and the aerodynamic load parameter k of this centrifugal force is aero The mechanical load parameter k of the centrifugal force is greater than or equal to 0.07 and less than or equal to 0.12. meca Greater than or equal to 11.5×10 6 tr 2 .m 2 .min -2 and less than or equal to 20.0×10 6 tr 2 .m 2 .min -2 On the other hand, the aerodynamic forces are higher than those of an engine with a conventional speed reducer, but are still acceptable. In particular, the fan shaft 20 of the propulsion system 1 (and indirectly the low-pressure shaft 11 ) is able to transmit the aerodynamic torque to the fan rotor 9 , for example, by means of a spline or bolt connection, without risking accidental damage or reducing the service life of the fan section 2 .

[0097] In order to reduce the mechanical load parameter k of the fan section 2 meca For example, the hub-to-tip ratio of the fan rotor 9 can be changed. The hub-to-tip ratio corresponds to the inner radius R of the fan rotor 9. i With outer radius R e The ratio between them can be changed by changing the diameter D of the fan rotor 9 (which is equal to the outer radius R e Twice, note that due to Figure 2 is a partial view, the diameter D is only partially visible) and / or by changing the inner radius R of the fan rotor 9 i However, increasing the diameter of the fan rotor 9 has the effect of increasing the centrifugal force applied to the fan blades 14. Thus, for example, by reducing the outer radius R of the fan rotor 9 e Or by increasing the inner radius R of the fan rotor 9 i to change the hub-to-tip ratio. For example, the hub-to-tip ratio of the fan rotor 9 may be between 0.22 and 0.32. In the case of a fixed-pitch fan rotor, the hub-to-tip ratio may be between 0.22 and 0.32. In the case of a variable-pitch fan rotor, for example, the hub-to-tip ratio may be between 0.24 and 0.32 to enable the integration of the pitch changing mechanism 15. Where appropriate, in order to achieve such a hub-to-tip ratio, the configuration of the fan shaft 20 and / or the bearings supporting the fan shaft 20 may be adjusted to facilitate the integration of the fan rotor 9 and various components of the propulsion system 1 arranged upstream of the reduction mechanism 19.

[0098] In order to obtain an aerodynamic load parameter k between 0.07 and 0.12 aero, the fan rotor 9 comprises at least twelve fan blades 14 and at most eighteen fan blades 14 , for example at least fourteen fan blades 14 and at most sixteen fan blades 14 .

[0099] The outer radius R of the fan rotor (9) e It can be between 40 inches (101.6 cm) and 92.5 inches (233.7 cm) (inclusive), for example between 60 inches (152.4 cm) and 85 inches (215.9 cm) (inclusive), for example about 478 inches (198.1 cm), which enables the propulsion system 1 to be integrated in a conventional manner, in particular under the wing of an aircraft.

[0100] Therefore, this outer radius R e Combined with a hub-to-tip ratio of between 0.22 and 0.32, this minimizes drag without compromising the fan rotor's mechanical loading, aerodynamic loading, or the fan's pressure ratio.

[0101] With the mechanical load parameter k as described above meca and the aerodynamic load parameter k aero The propulsion system 1 can have a high bypass ratio, even when the thrust rating of the propulsion system is moderate. Generally, the propulsion system 1 can be 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), and a bypass ratio greater than or equal to 10, particularly greater than or equal to 40, such as between 40 and 80 inclusive. Thus, the propulsion system 1 can have a high bypass ratio while still being able to be integrated under the wing of the aircraft 100.

[0102] In the case of one or more unducted fan rotors 9 , the peripheral speed at the tips 21 of the blades 14 of one or more fan rotors 9 may also be between 210 m / s and 260 m / s.

[0103] The reduction ratio of the reduction mechanism 19 may be between 5.0 and 11.0. This reduction ratio enables a fan section 2 to be obtained in which the fan rotor 9 rotates at a rotation speed ω such that the aerodynamic load parameter is greater than or equal to 0.07 and less than or equal to 0.12, and the mechanical load parameter is greater than or equal to 11.5×10 6 tr 2 .m 2 .min -2 and less than or equal to 20.0×10 6 tr2 .m 2 .min -2 .

[0104] Having an aerodynamic load parameter k greater than or equal to 0.07 and less than or equal to 0.12 aero and greater than or equal to 11.5×10 6 tr 2 .m 2 .min -2 and less than or equal to 20.0×10 6 tr 2 .m 2 .min -2 Mechanical load parameter k meca The twin-body propulsion system 1 may in particular include a two-stage high-pressure turbine 7, a high-pressure compressor 5 including at least eight stages and at most eleven stages, a low-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.

[0105] Comparison example:

[0106] The engine 1 is a twin-body propulsion system comprising an unducted fan section 2 corresponding to the current technical standard (at the filing date of this application) to which we are seeking to improve.

[0107] The engine 2 is a twin-body propulsion system 1 including an unducted fan section 2 according to the teachings of the present application having an aerodynamic load parameter k equal to 0.075. aero The sum is 14.6×10 6 tr 2 .m 2 .min -2 Mechanical load parameter k meca .

[0108]

[0109]

[0110] The aerodynamic load parameter of engine 1 is less than 0.07, and the mechanical load parameter of engine 1 is less than 11.5×10 6 tr 2 .m 2 .min -2 The engine 1 has low aerodynamic loads and low mechanical loads, which improves fan efficiency at the expense of fan size and mass. The size of the fan has a significant impact on the installation of the engine on the aircraft.

[0111] The integration of engine 2 into the aircraft is simplified compared to engine 1 without reducing the energy performance of engine 2. By increasing the aerodynamic load and the fan pressure ratio, the fan diameter can be reduced while maintaining thrust.

[0112] This is facilitated by an increase in mechanical load, which increases the fan rotor speed.

[0113] Compared to engine 1, the fan section 2 of engine 2 is subjected to slightly greater mechanical and aerodynamic stresses, in particular during takeoff, while remaining within an acceptable range for the engine dimensions, enabling the fan shaft 10 to transmit the aerodynamic torque to the fan rotor 9. On the other hand, due in particular to the reduction in diameter and bypass ratio of engine 2, engine 2 is more compact and has a lower mass than engine 1, without reducing the efficiency of engine 2. In this comparative example, the reduction in mass of the fan section 2 is estimated to be approximately 20%. Insofar as the fan section represents one-third of the mass of the propulsion system 1, this is equivalent to reducing the mass of the propulsion system 1 by approximately 7%, with attendant effects on the aircraft (cantilever mass, reduction in diameter of the fan rotor 9, etc.). As a result, engine 2 can be more easily mounted on aircraft 100, having the same inlet temperature and the same fan thrust as the low-pressure turbine 8.

[0114] In order to move from the (reference) engine 1 to the (disclosed) engine 2, the fan diameter D and the bypass ratio BPR have been reduced, thereby improving the integration of the engine 2. On the other hand, the pressure ratio of the fan section of the engine 2 has been slightly increased (while remaining less than 1.2) to maintain equivalent thrust. In addition, the overall compression ratio has been increased without increasing the inlet temperature of the high-pressure turbine 7, which improves the efficiency of the body without increasing the thermal load on the low-pressure turbine 8. Finally, since the temperature of the low-pressure turbine 8 remains stable, the mechanical load (N1) of the low-pressure turbine 8 can be increased. 2 S), to reduce the number of stages of the low-pressure turbine 8.

[0115] By reducing the fan blades, the value of the aerodynamic load parameter can be increased with limited impact on fan efficiency. Reducing the elongation of the fan blades is generally considered to be detrimental to fan efficiency and blade vibration conditions relative to the first torsional mode.

[0116] However, improved aerodynamics can further minimize the impact on fan efficiency. Reducing blade height and using three-dimensional woven composites with appropriate weave patterns or incorporating stiffer fibers can restore vibration conditions relative to the first torsional mode.

[0117] The use of fan blades comprising a three-dimensional braided composite material with stiffer fibers also makes it possible to withstand increased centrifugal mechanical stresses, particularly in the event of bird ingestion. The use of bearing steel or ceramic rolling elements with increased mechanical properties makes it possible to achieve the same hub-to-hub ratio on Engine 2 as on Engine 1, despite the increased mechanical load parameters. The use of titanium alloys with improved mechanical properties for the hub contributes to achieving the target hub-to-hub ratio.

[0118] The overall compression ratio of the engine 2 and the high-pressure turbine inlet temperature will also increase to improve the thermal efficiency of the engine.

Claims

1. An aviation propulsion system (1), comprising: - a drive shaft (11) rotatable about an axis of rotation; - fan shaft (20); - a fan section (2), said fan section comprising an unducted fan rotor (9) rotated by said fan shaft (20), said fan rotor (9) comprising a plurality of blades (14); - a speed reduction mechanism (19) coupling the drive shaft (11) and the fan shaft (20) so as to drive the fan shaft (20) at a rotation speed lower than that of the drive shaft (11); The propulsion system (1) is configured such that the fan rotor (9) has an aerodynamic load parameter greater than or equal to 0.07 and less than or equal to 0.12, and a rotational speed greater than or equal to 11.5×10 6 tr.m 2 .min -1 and less than or equal to 20.0×10 6 tr.m 2 .min -1 Mechanical load parameters, The aerodynamic load and mechanical load are defined by the following formulas: k meca =ω 2 ×S Where: k aero is the aerodynamic load parameter; k meca is the mechanical load parameter; FN is the thrust generated by the propulsion system (1) when the propulsion system (1) is at rest, take-off speed, standard atmospheric pressure and sea level conditions, expressed in Newtons (N); BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is at rest, takeoff speed, standard atmospheric pressure and sea level conditions; n is the number of blades (14) in the fan rotor (9); F=39.7N -1.35 .(tr / min) -2 .m -5 ; R e is the outer radius of the fan rotor (9) and corresponds to the distance between the axis of rotation (X) and the intersection between the tip (21) and the leading edge (22) of the blades (14) of the fan rotor (9), expressed in meters (m); R i is the inner radius of the fan rotor and corresponds to the distance between the axis of rotation (X) and the intersection of the leading edges of the blades (14) of the fan rotor (9) and the surface of the fan rotor (9) which delimits the flow stream in the fan rotor (9) radially inside, this distance being expressed in meters (m); ω is the limit speed of the fan rotor (9), expressed in revolutions per minute (tr / min); and S is the surface area of ​​the flow stream at the inlet of the fan rotor (9), which is equal to π×(R e 2 -R i 2 ), in square meters (m 2 )express.

2. The propulsion system (1) according to claim 1, wherein: The outer radius (Re) of the fan rotor (9) is between 40 inches (101.6 cm) and 92.5 inches (233.7 cm) inclusive, for example, between 60 inches (152.4 cm) and 85 inches (215.9 cm) inclusive, for example, about 478 inches (198.1 cm).

3. The propulsion system (1) according to claim 1 or 2, wherein: The fan rotor (9) has a hub-to-tip ratio between 0.22 and 0.

32.

4. The propulsion system (1) according to any one of claims 1 to 3, wherein: The fan rotor (9) comprises at least twelve fan blades (14) and at most eighteen fan blades (14), for example at least fourteen fan blades (14) and at most sixteen fan blades (14).

5. The propulsion system (1) according to any one of claims 1 to 4, wherein: The reduction ratio of the reduction mechanism (19) is greater than or equal to 5.0 and less than or equal to 11.

0.

6. The propulsion system (1) according to any one of claims 1 to 5, wherein: The bypass ratio of the propulsion system (1) is greater than or equal to 40, for example, between 40 and 80 and including 40 and 80.

7. The propulsion system (1) according to any one of claims 1 to 6, wherein: When the propulsion system (1) is at rest, take-off speed, standard atmospheric pressure and sea level conditions, the peripheral speed at the tips (21) of the blades (14) of the fan rotor (9) is between 210 m / s and 260 m / s.

8. The propulsion system (1) according to any one of claims 1 to 7, wherein: The fan section (2) also has a fan compression ratio, which corresponds to a pressure ratio between an outlet of the fan rotor (9) and an inlet of the fan rotor (9), which is less than or equal to 1.45, for example less than or equal to 1.

30.

9. The propulsion system (1) according to any one of claims 1 to 8, wherein: The propulsion system (1) is sized so that the thrust of the propulsion system is between 18,000 lbf (80,068 N) and 51,000 lbf (222,411 N) when the propulsion system (1) is at rest, takeoff speed, and standard atmospheric pressure conditions.

10. The propulsion system (1) according to any one of claims 1 to 9, further comprising a drive turbine (7) and a compressor (4) directly connected via the drive shaft (11), wherein: The driving turbine (7) comprises at least three stages and at most five stages.

11. The propulsion system (1) according to claim 10, wherein: The compressor (4) comprises at least two stages and at most four stages.

12. The propulsion system (1) according to claim 10 or 11, further comprising a high-pressure turbine (8) and a high-pressure compressor (5) connected via a high-pressure shaft (10), wherein the high-pressure shaft (10) rotates faster than the drive shaft (11), and the high-pressure turbine (8) is two-stage.

13. The propulsion system (1) according to claim 12, wherein: The high-pressure compressor (5) comprises at least eight stages and at most eleven stages.

14. An aircraft (100) comprising at least one propulsion system (1) according to any one of claims 1 to 13, said propulsion system being fixed to the aircraft via a mast.

15. A method for determining the size of a propulsion system (1), the propulsion system comprising a speed reduction mechanism (19) coupling a drive shaft (11) and an unducted fan rotor (9) so as to drive the fan rotor (9) at a speed lower than the speed of the drive shaft (11); wherein: The fan rotor (9) has an aerodynamic load parameter greater than or equal to 0.07 and less than or equal to 0.12, and a load factor greater than or equal to 11.5×10 6 tr.m 2 .min -1 and less than or equal to 20.0×10 6 tr.m 2 .min -1 Mechanical load parameters, The aerodynamic load and mechanical load are defined by the following formulas: k meca =ω 2 ×S Where: k aero is the aerodynamic load parameter; k meca is the mechanical load parameter; FN is the thrust generated by the propulsion system (1) when the propulsion system (1) is at rest, take-off speed, standard atmospheric pressure and sea level conditions, expressed in Newtons (N); BPR is the bypass ratio of the propulsion system and is measured when the propulsion system is at rest, takeoff speed, standard atmospheric pressure and sea level conditions; n is the number of blades (14) in the fan rotor (9); F=39.7N -1.35 .(tr / min) -2 .m -5 ; R e is the outer radius of the fan rotor (9) and corresponds to the distance between the axis of rotation (X) and the intersection between the tip (21) and the leading edge (22) of the blades (14) of the fan rotor (9), expressed in meters (m); R i is the inner radius of the fan rotor and corresponds to the distance between the axis of rotation (X) and the intersection of the leading edges of the blades (14) of the fan rotor (9) and the surface of the fan rotor (9) which delimits the flow stream in the fan rotor (9) radially inside, this distance being expressed in metres (m); and ω is the limit speed of the fan rotor (9), expressed in revolutions per minute (tr / min); and S is the surface area of ​​the flow stream at the inlet of the fan rotor (9), which is equal to π×(R e 2 -R i 2 ), in square meters (m 2 )express.

16. A method for manufacturing an aviation propulsion system, comprising the steps of: - sizing the aerospace propulsion system according to the method of claim 15; and - Manufacturing of said aviation propulsion system.

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