Aero-propulsion system including an optimized fan section

Abstract

This disclosure relates to a fan section with a performance coefficient between 1.05 and 1.3, the performance coefficient being defined as follows: (Cp) where: Cp is the performance coefficient; FPR is the pressure ratio of the fan section (2); Cs is the solidity of the fan rotor (9); and U is the peripheral velocity at the blade tip of the fan rotor (9); the pressure ratio and peripheral velocity are measured at cruising speed.

Description

Technical Field

[0001] This application relates generally to the field of propulsion systems, and more specifically to aerospace propulsion systems with high or extremely high bypass ratios. Background Technology

[0002] A propulsion system typically includes, from upstream to downstream, a fan section, a compressor section, a combustion chamber, and a turbine section along the direction of gas flow. 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 rotatably via a high-pressure turbine through a high-pressure shaft. The fan and the low-pressure compressor (where appropriate) are driven rotatably via a low-pressure turbine through a low-pressure shaft.

[0003] Technological research has significantly improved the environmental performance of aircraft. The applicant has considered all factors at all design and development stages to obtain components and aviation products that consume less energy, are more environmentally friendly, and can be integrated and used in civil aviation with a moderate environmental impact, with the aim of improving the energy efficiency of aircraft.

[0004] Therefore, to improve the propulsion efficiency of the propulsion system and reduce the unit fuel consumption and noise emitted by the fan section, propulsion systems with a high bypass ratio (BPR) (equivalent to the ratio of secondary airflow to primary airflow) have been proposed. To achieve such a bypass ratio, the fan section can be separated from the low-pressure turbine, allowing for independent optimization of the respective rotational speeds of the fan section and the low-pressure turbine. Typically, this separation is achieved using a reduction gear located between the upstream end of the low-pressure shaft and the rotor of the fan section. The fan section rotor is then driven by the low-pressure shaft via the reduction gear at a speed lower than the low-pressure shaft's rotational speed.

[0005] Improving the propulsion efficiency of the system may involve determining the dimensions of the fan section. Specifically, due to the wide diameter of the fan section (especially to achieve a high bypass ratio and low fan pressure ratio), the fan section accounts for a considerable portion of the propulsion system in terms of mass and fuel consumption per unit. At the same time, the fan section generates a significant portion of the thrust of the propulsion system. Summary of the Invention

[0006] One objective of this application is to optimize the fan section of the propulsion system to make the fan section more efficient without causing excessive losses in mass and unit fuel consumption of the propulsion system.

[0007] Therefore, according to a first aspect, a fan section for an aircraft propulsion system is provided, the fan section including a fan rotor with a performance coefficient greater than or equal to 1.05 and less than or equal to 1.3, wherein the performance coefficient is defined as follows:

[0008] Where: Cp is the performance coefficient, with units of seconds per square meter (s / m)²; FPR is the pressure ratio of the fan section; Cs is the solidity of the fan rotor; and U is the peripheral velocity at the tip of the fan rotor blades, measured in meters per second (m / s). Pressure ratio and ambient velocity are measured at cruising speed.

[0009] According to the first aspect, some preferred but non-limiting features of the fan section are the following features, used alone or in combination: - The fan pressure ratio is greater than or equal to 1.1 and less than or equal to 1.45; - The solidity is greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.1; - The surrounding speed is greater than or equal to 260 m / s and less than or equal to 400 m / s, for example, greater than or equal to 270 m / s and less than or equal to 350 m / s; - The extension ratio of a fan blade is the ratio of the height of the fan blade to the chord length at the tip of the fan blade. The extension ratio of a fan blade is greater than or equal to 2.4 and less than or equal to 3.0. - The diameter of the fan rotor is greater than or equal to 177.8 cm and less than or equal to 304.8 cm, for example, less than or equal to 304.8 cm, such as approximately 228.6 cm; and / or - The ratio of the hub to the tip of the fan section is greater than or equal to 0.22 and less than or equal to 0.32, for example greater than or equal to 0.235 and less than or equal to 0.30, or again for example less than or equal to 0.27.

[0010] According to the second aspect, an aviation propulsion system is provided, comprising: - Drive shaft, which is capable of rotating about a rotation axis; - Fan shaft; -According to the fan section of the first aspect, the fan rotor is driven to rotate by the fan shaft; and - A reduction gear connects the drive shaft and the fan shaft, enabling the fan shaft to be driven at a speed lower than that of the drive shaft.

[0011] Optionally, the aviation propulsion system according to the second aspect may also include a fan housing surrounding the fan blades.

[0012] According to the second aspect, the bypass ratio of the propulsion system can also be greater than or equal to 10, for example, between 10 and 35 (inclusive), for example, between 10 and 18 (inclusive).

[0013] According to a third aspect, an aircraft is provided, the aircraft including at least one propulsion system according to a second aspect, the propulsion system being attached to the aircraft via a mast.

[0014] According to a fourth aspect, a method is provided for determining the dimensions of a fan section of an aircraft propulsion system. This method includes determining the pressure ratio of the fan section, the solidity of the fan section's rotor, and the peripheral velocity of the rotor, such that the performance coefficient of the fan section is greater than or equal to 1.05 and less than or equal to 1.3, wherein the performance coefficient is defined as follows:

[0015] Where: Cp is the performance coefficient; FPR is the pressure ratio of the fan section; Cs is the solidity of the fan rotor; and U is the peripheral velocity at the blade tip of the fan rotor; the pressure ratio and peripheral velocity are measured at cruising speed.

[0016] Optionally, the dimension determination method according to the fourth aspect further includes the step of determining the dimensions of the fan blades, such that the extension ratio of the fan blades is greater than or equal to 2.4 and less than or equal to 3.0, wherein the extension ratio is the ratio of the height of the fan blades to the chord length at the tip of the fan blades.

[0017] According to the fifth aspect, a method for manufacturing a fan section is proposed, the method comprising the following steps: - Determine the dimensions of the fan section according to the dimensional determination method in the fourth aspect; and - Manufacture fan sections of such specific dimensions. Attached Figure Description

[0018] Other features, objects, and advantages of the invention will become apparent from the following description, which is merely illustrative and non-limiting and should be read with reference to the accompanying drawings, in which: Figure 1 This is a schematic partial cross-sectional view of an example of a propulsion system according to the first embodiment; Figure 2a This is a schematic partial cross-sectional view of an example fan rotor of a propulsion system according to an embodiment, the cross-section being located in a plane passing through the upstream intersection between the tips and leading edges of two adjacent blades; Figure 2b This is a perspective view of an example fan rotor of a propulsion system according to an embodiment; Figure 3 This is a schematic cross-sectional view of an example of a first variant of the reduction gear mechanism; Figure 4 This is a schematic cross-sectional view of an example of a second variant of a planetary deceleration mechanism; Figure 5 This is an example of an aircraft that may include at least one propulsion system according to an embodiment; Figure 6 This is a flowchart illustrating an example of steps in a method for determining or manufacturing dimensions according to an embodiment.

[0019] In all the figures, similar elements have the same reference numerals. Detailed Implementation

[0020] The propulsion system 1 has a main direction extending along the longitudinal axis X. When the propulsion system 1 is in operation, it includes, from upstream to downstream, a fan section 2 and a main body 3 (commonly referred to as a "gas generator") in the direction of gas flow through the propulsion system 1. The main body 3 includes compressor sections 4 and 5, a combustion chamber 6, and turbine sections 7 and 8. Here, the propulsion system 1 is an aerospace propulsion system 1, which is configured to be attached to the aircraft 100 via a pylon (or mast).

[0021] Compressor sections 4 and 5 include a series of stages, each stage including movable impellers (rotors) 4a and 5a that rotate in front of fixed impellers (stators) 4b and 5b. Turbine sections 7 and 8 also include a series of stages, each stage including fixed impellers (stators) 7b and 8b, and movable impellers (rotors) 7a and 8a that rotate behind the fixed impellers (stators) 7b and 8b.

[0022] In this application, the axial direction is the direction of the longitudinal axis X, which corresponds to the rotation of the gas generator shaft, and the radial direction is the direction perpendicular to and through the axis X. Furthermore, the circumferential (or lateral or tangential) direction is the direction perpendicular to the longitudinal axis X but does not pass through it. Unless otherwise stated, "inner" and "outer" are used with reference to the radial direction, such that the inner portion or inner surface of the element is closer to the axis X than the outer portion or outer surface of the same element.

[0023] During operation, the airflow F entering the propulsion system 1 is divided into a primary airflow F1 and a secondary airflow F2, which flow from upstream to downstream through the propulsion system 1.

[0024] The secondary airflow F2 (also known as the "bypass airflow") flows around the main body 3. The secondary airflow F2 enables the cooling of the periphery of the main body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0025] The main airflow F1 flows through the main air path within the main body 3, passing sequentially through compressor sections 4 and 5, combustion chamber 6, and turbine sections 7 and 8. In combustion chamber 6, the main airflow F1 mixes with fuel to act as an oxidizer. The main airflow F1, by passing through turbine sections 7 and 8 which receive energy from combustion chamber 6, causes the rotors of turbine sections 7 and 8 to rotate, which in turn drives the rotors of compressor sections 4 and 5 and the rotor portion 9 of fan section 2.

[0026] In the 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 the high-pressure compressor 5 is driven rotatably by the rotor of the high-pressure turbine 7 via a high-pressure shaft 10. The rotor of the low-pressure compressor 4 and the rotor portion 9 of the fan section 2 are driven rotatably by the rotor of the low-pressure turbine 8 via a low-pressure shaft 11. Therefore, the main body 3 includes a high-pressure main body and a low-pressure main body. The high-pressure main body includes the high-pressure compressor 5, the high-pressure turbine 7, and the high-pressure shaft 10, while the low-pressure main body includes the fan section 2, the low-pressure compressor 4, the high-pressure turbine 8, and the low-pressure shaft 11. The rotational speed of the high-pressure main body is greater than that of the low-pressure main body. In the three-body propulsion system 1, turbine sections 7 and 8 also include an intermediate turbine located between the high-pressure turbine 7 and the low-pressure turbine 8 and configured to drive the rotor of the low-pressure compressor 4 via an 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.

[0027] The low-pressure shaft 11 is typically housed within and coaxial with the high-pressure shaft 10 for a section of its length. 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. In a variant, the low-pressure shaft 11 and the high-pressure shaft 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.

[0028] The fan section 2 includes at least a fan rotor 9, which is adapted to be rotatably driven relative to the stator portion of the propulsion system 1 by 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 may be fixed relative to the hub 13.

[0029] The fan section 2 may also include a fan stator 16 or a straightener, which includes blades 17 mounted on the hub of the fan stator 16 and has the function of straightening the secondary airflow F2 flowing at the outlet of the fan rotor 9. The blades 17 of the fan stator 18 may be fixed relative to the hub or have a variable setting.

[0030] To improve the propulsion efficiency of propulsion system 1 and reduce its unit fuel consumption and the noise emitted by fan section 2, propulsion system 1 has a high bypass ratio. The term "high bypass ratio" is understood herein to mean a bypass ratio greater than or equal to 10, for example, between 10 and 80 (inclusive). To calculate the bypass ratio, the mass flow rates of the secondary airflow F2 and the primary airflow F1 are measured when propulsion system 1 is stationary, not installed, at standard atmospheric pressure (as defined by the International Civil Aviation Organization (OACI) manual, document 7488 / 3, 3rd edition) and at takeoff speed below sea level. The term "not installed" is understood herein to mean that the measurement is performed when propulsion system 1 is on a test stand (and not installed on aircraft 100), as this measurement is easier to perform.

[0031] Note that in this application, certain parameters were determined under cruise conditions, i.e., at an altitude of 10,668 m (35,000 feet), a flight speed of Mach 0.8, and under International Standard Atmosphere (ISA) conditions as defined in Annex 1985 of ISO 2533 / 1975. Furthermore, when the propulsion system 1 is cooled, i.e., when the propulsion system 1 has been stopped for a sufficient period of time to allow the components of the propulsion system to reach ambient temperature, distances (length, radius, diameter, chord length, etc.) are measured at ambient temperature (approximately 20°C). It should be understood that these dimensions change very little relative to the conditions at takeoff speed when the propulsion system 1 is at its takeoff speed.

[0032] A reduction gear 19 is used to separate the fan rotor 9 from the low-pressure shaft 11, allowing for independent optimization of the respective rotational speeds of the fan rotor 9 and the low-pressure shaft 11. The reduction gear 19 is positioned between the upstream end of the low-pressure shaft 11 and the fan rotor 9. In this case, the propulsion system 1 also includes an additional shaft (i.e., the so-called 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. Thus, the fan rotor 9 is driven by the low-pressure shaft 11 via the reduction gear 19 and the fan shaft 20 at a speed lower than that of the low-pressure turbine 8.

[0033] This separation allows for a reduction in the speed and pressure ratio of the fan rotor 9 and an increase in the power extracted by the high-pressure turbine 8. Specifically, the overall efficiency of the propulsion system primarily depends on 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 the propulsion system 1 with a high bypass ratio, most of the flow generating propulsion is formed by the secondary airflow F2 of the propulsion system 1, and the kinetic energy of the secondary airflow F2 is mainly affected by the compression it undergoes as it passes through the fan section 2. Therefore, the propulsion efficiency is related to the pressure ratio of the fan section 2: the lower the pressure ratio of the fan section 2, the higher the propulsion efficiency.

[0034] When propulsion system 1 is stationary, uninstalled, at standard atmospheric pressure (as defined by the International Civil Aviation Organization (OACI) manual, document 7488 / 3, 3rd edition) and at takeoff speed below sea level, propulsion system 1 is configured to provide thrust between 18,000 lbf (80,068 N) and 51,000 lbf (222,411 N), for example between 20,000 lbf (88,964 N) and 35,000 lbf (155,688 N).

[0035] The fan section 2 may include a fan housing 12, in which a fan rotor 9 is housed.

[0036] The fan rotor 9 extends upstream of the fan stator 16. The blades of the fan stator are typically referred to as outlet guide vanes (OGVs) and have a fixed configuration relative to the hub of the fan stator. Furthermore, the bypass ratio of the propulsion system 1 is, for example, greater than or equal to 10, for example, between 10 and 35 (inclusive), for example, between 10 and 18 (inclusive).

[0037] Each fan blade 14 has a leading edge 14a and a trailing edge 14b (see, for example, see...). Figure 2a and Figure 2b The leading edge 14a is configured to extend towards the gas flow entering the fan rotor 9. The leading edge 14a corresponds to a front portion with an aerodynamic profile that faces the airflow and divides it into a pressure surface flow and a suction surface flow. Simultaneously, the trailing edge 28b corresponds to a rear portion with an aerodynamic profile where the pressure surface flow and the suction surface flow intersect. Note that when the blade 14 includes leading and / or trailing edge sealing rings, the leading edge 14a (or trailing edge 14b) of the blade 14 corresponds to a reconstruction of the sealing ring profile as the leading edge (or a reconstruction of the sealing ring profile as the trailing edge), and the function of this leading edge (or trailing edge) is to divide the flow into a pressure surface flow and a suction surface flow (or to cause the flows to intersect).

[0038] In addition, the fan rotor 9 includes a series of platforms, each extending between two adjacent blades 14, and is configured to radially define the airflow F passing through the rotor 9 internally.

[0039] Furthermore, the fan blade 14 has a chord length c1 at its tip. The chord length c1 at the blade tip is equivalent to the line segment connecting the upstream intersection point P between the leading edge 14a of the blade 14 and the tip 21, and the downstream intersection point between the trailing edge 14b of the blade 14 and the tip 21.

[0040] The fan rotor 9 comprises 16 to 22 blades 14.

[0041] The reduction mechanism 19 may include a reduction mechanism 19 having planetary gears, such as a single-stage or two-stage reduction mechanism of the "planetary" or "star" type as known to those skilled in the art. In a first variation, the reduction mechanism 19 may be of the "star" type. Figure 3 The system includes a sun gear 19a (the inlet of the reduction gear 19), a ring gear 19b (the outlet of the reduction gear 19), and a series of planetary gears 19c. The sun gear 19a is centered on the rotational axis X of the reduction gear 19 (generally collinear with the longitudinal axis X) and is configured to be rotatably driven by the low-pressure shaft 11. The ring gear 19b is coaxial with the sun gear 19a and is configured to rotatably drive the fan shaft 20 about the rotational axis X. A series of planetary gears 19c are circumferentially distributed between the sun gear 19a and the ring gear 19b about the rotational axis X. Each planetary gear 19c meshes internally with the sun gear 19a and externally with the ring gear 19b. This series of planetary gears 19c is mounted on a planet carrier 19d, which is fixed relative to the stator portion 19e of the propulsion system 1, for example, relative to the housing of the compressor sections 4 and 5. In a second variation, the reduction gear 19 may be planetary (…). Figure 4 In this case, the gear ring 19b is fixedly mounted on the stator portion 19e of the propulsion system 1, and the fan shaft 20 is rotatably driven by the planet carrier 19d (therefore, the planet carrier 19d can rotate relative to the stator portion 19e of the propulsion system 1, for example, relative to the housing of the compressor sections 4 and 5).

[0042] Regardless of the configuration of the reduction mechanism 19, the diameters of the gear ring 19b and the planetary carrier 19d are larger than the diameter of the sun gear 19a, so that the speed of the fan rotor 9 is less than the speed of the low-pressure shaft 11.

[0043] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11, for example, greater than or equal to 2.7 and less than or equal to 6.0, for example, about 3.0.

[0044] The dual-body propulsion system 1 may specifically include two-stage high-pressure turbines 7, at least eight and at most eleven-stage high-pressure compressors 5, at least three and at most five-stage low-pressure turbines 8, and at least two and at most four-stage low-pressure compressors 5.

[0045] When propulsion system 1 is stationary, uninstalled, at standard atmospheric pressure (as defined by the International Civil Aviation Organization (OACI) manual, document 7488 / 3, 3rd edition) and below sea level, the limiting speed of the low-pressure shaft 11 is between 8500 rpm and 12000 rpm, for example, between 9000 rpm and 11000 rpm. This limiting speed 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-E740 (or according to US certification regulation 14-CFR Part 33.87)). The limiting speed corresponds to the maximum rotational speed of propulsion system 1 when it is healthy. Therefore, the low-pressure shaft 11 is likely to reach the limiting speed under flight conditions. This limiting speed is part of the data declared in the type certificate datasheet. Specifically, this rotational speed is commonly used as a reference speed for dimensional determination and manufacturing of propulsion system 1, and is used for certain certification tests (such as blade loss or rotor integrity tests, typically CS-E-800 certification - bird strike and ingestion).

[0046] To optimize the performance of propulsion system 1, the performance coefficient of the fan section is between 1.05 and 1.3, whereby the performance coefficient is defined as follows:

[0047] Where: Cp is the performance coefficient, with units of (seconds / meter)² (s / m)²; FPR is the pressure ratio of fan section 2 (dimensionless). Cs is the real number (dimensionless) of the fan rotor 9; and U is the peripheral velocity at the tip of the fan rotor blades, measured in meters per second (m / s).

[0048] Note that here, the pressure ratio FPR and the peripheral speed U of fan section 2 are measured at cruise speed, because cruise speed is the flight stage where the maximum efficiency of fan rotor 9 is desired.

[0049] The solidity Cs is equal to the ratio of the chord length c1 at the blade tip to the inter-blade pitch 23. The inter-blade pitch 23 corresponds to the angular distance between the upstream intersection points P of two adjacent blades 14; therefore, the inter-blade pitch 23 is equal to the outer radius R of the fan rotor 9. e (Half the diameter) multiplied by the angle between the first line D1 and the second line D2, the first line being contained in a plane perpendicular to the axis X, the plane intersecting at the upstream point P of the first blade 14 (see...). Figure 2a and Figure 2b The second line begins at the point P, an upstream intersection of the first blade 14 and the axis X, and is contained in a plane perpendicular to the axis X. (See [reference]). Figure 2bSince realism is a ratio of distance, it is measured when propulsion system 1 is cooled (under the aforementioned conditions). For example, realism Cs can be greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.1.

[0050] The pressure ratio FPR of fan section 2 is equivalent to the ratio of the average pressure at the outlet of fan stator 17 to the average pressure at the inlet of fan rotor 9. For example, the pressure ratio is greater than or equal to 1.1 and less than or equal to 1.45. Here, the average pressure is measured along the airflow path (from the surface that defines the airflow path at the inlet of fan rotor 9 radially inward to the fan housing 12).

[0051] For example, the peripheral velocity U is greater than or equal to 260 m / s and less than or equal to 400 m / s, such as greater than or equal to 270 m / s and less than or equal to 350 m / s.

[0052] The dimensions of fan section 2 are determined such that its coefficient of performance (Cp) is between 1.05 and 1.3, maximizing its performance at cruise speed while maintaining acceptable maneuverability (primarily during takeoff). Specifically, the Cp takes into account the aerodynamic work that the fan rotor 9 must perform (via the fan rotor's pressure ratio FPR) and the physical material of the blades 14 that perform this work (via solidity Cs). A higher Cp allows the fan rotor 9 to deliver a greater amount of aerodynamic work with smaller blades 14. Therefore, the efficiency of the fan rotor 9 is improved, while its size and mass are reduced. The Cp is preferably less than or equal to 1.3 to avoid any surge and maintain sufficient mechanical strength in the event of bird strikes or inhalation.

[0053] For example, fan section 2 may have a pressure ratio (FPR) of 1.35, a solidity of 0.9, a peripheral velocity of 305 m / s, and a performance coefficient of 1.3. In another example, fan section 2 may have a pressure ratio (FPR) of 1.38, a solidity of 1.05, a peripheral velocity of 315 m / s, and a performance coefficient of 1.05. In yet another example, fan section 2 may have a pressure ratio (FPR) of 1.3, a solidity of 0.95, a peripheral velocity of 295 m / s, and a performance coefficient of 1.3.

[0054] Preferably, the blades 14 of the fan rotor 9 also have an elongation ratio greater than or equal to 2.4 and less than or equal to 3.0. The elongation ratio is the ratio of the height h of the fan blade to the chord length c1 of the fan blade at its tip. When the performance coefficient of the fan section is between 1.05 and 1.3, and the elongation ratio of the blades 14 is between 2.4 and 3.0, the thickness of the fan blades 14 can be reduced while meeting current certification requirements regarding bird strikes (typically CS-E-800 certification - bird strike and bird ingestion). The fan blades 14 are more "conical," which further reduces the mass of the fan section 2, thus reducing the unit fuel consumption of the propulsion system 1, while ensuring the required aerodynamic work and avoiding the risk of surge. Furthermore, the overall length of the fan nacelle can be reduced.

[0055] Here, the height h of the blade 14 is measured in a plane perpendicular to the axis X, between the upstream intersection point P (between the tip 21 and leading edge 14a of the blade 14 of the fan rotor 9) and the intersection point between the blade and the platform in a plane perpendicular to the axis X. Figure 1 ).

[0056] The diameter D of the fan rotor can be between 70 inches (177.8 cm) and 185 inches (469.9 cm) (inclusive). For example, a diameter D between 70 inches (177.8 cm) and 120 inches (304.8 cm) (inclusive), such as approximately 90 inches (228.6 cm), allows for the conventional integration of the propulsion system 1, particularly under the wing of the aircraft 100. The diameter of the fan rotor 9 is measured in a plane perpendicular to the axis of rotation X, at the upstream intersection point P (between the tip 21 of the blade 14 of the fan rotor 9 and the leading edge 14a). Note that due to... Figure 1 If it is a partial view, then only part of the diameter D is visible.

[0057] The number of blades 16 in the fan stator 17 depends on the acoustic criteria defined for the propulsion system 1 and is at least equal to the number of blades 14 in the fan rotor 9. In a variation of the embodiment, the number of blades 16 in the fan stator 17 is at least 30 and at most 52, for example, exactly 48 when the fan rotor includes twenty-two blades 14.

[0058] Furthermore, for example, the fan rotor 9 has a hub-to-tip ratio between 0.22 and 0.32. With a fixed setting for the fan rotor 9, the hub-to-tip ratio can be between 0.22 and 0.30, for example, between 0.235 and 0.27. The hub-to-tip ratio is the ratio of the inner radius R of the fan rotor 9. i With outer radius R e The ratio of the inner radius R. iIt is the distance between the rotation axis X and the intersection point (corresponding to the connection point of the leading edge 22 and the aerodynamic surface of the fan rotor 9 platform), which is located between the leading edge 22 and the surface that defines the airflow path at the inlet of the fan rotor 9 on the radially inner side. Outer radius R e It is equal to half the fan diameter D. The lower the hub-to-tip ratio, the better the performance of the fan rotor 9. However, a decrease in the hub-to-tip ratio of the fan rotor 9 leads to an increase in the mechanical load on the hub 13 of the fan rotor 9.

[0059] Example: Example 1 In a first example of an engine, the fan section 2 has a performance coefficient between 1.05 and 1.3, comprising 22 fan blades 14, a peripheral velocity U at cruising speed of 315 m / s, a pressure ratio FPR at cruising speed of 1.4, and a solidity of 1.0. The performance coefficient of this fan section 2 is 1.1 (s / m)². Furthermore, this fan section 2 includes 48 stator blades 16. The fan blades 14 have a chord length c1 of 32.6 cm at the tip, a height of 84.6 cm, and an extension ratio of 2.8. Additionally, the fan rotor 9 has a diameter of 229 cm and a hub-to-tip ratio of 0.26. The fan blades are made of a composite material comprising embedded fiber reinforcements (including carbon fibers with a Young's modulus greater than 250 GPa) in an epoxy resin matrix. The fan section is ducted, and the fan blades 14 have a variable setting.

[0060] When propulsion system 1 is stationary, uninstalled, at standard atmospheric pressure (as defined by the International Civil Aviation Organization (OACI) manual, document 7488 / 3, 3rd edition) and takeoff speed below sea level, propulsion system 1 is configured to provide 164 kN of thrust.

[0061] This fan section 2 can be integrated into a propulsion system 1, which has a bypass ratio of 14 at cruising speed and includes a reduction gear 19 with a reduction ratio of 3.2. Furthermore, the propulsion system 1 may include two low-pressure compressor stages (4 stages), ten high-pressure compressor stages (5 stages), two high-pressure turbine stages (7 stages), and five low-pressure turbine stages (8 stages).

[0062] Compared to fan sections with a performance coefficient of less than 1.05 (s / m)² (e.g., equal to 0.7), the fan section of the above-described engine example is able to achieve a better level of efficiency with lower mass and volume, while maintaining an acceptable level of operability with respect to certification.

[0063] Example 2 Engine 1 is a dual-body propulsion system, which includes a ducted fan section 2 corresponding to the current technical standard (the filing date of this application), and it is desired to improve upon the ducted fan section 2.

[0064] The engine 2 is a dual-body propulsion system 1, which includes a ducted fan section 2 with a performance coefficient between 1.05 and 1.3 according to the teachings of this application.

[0065]

[0066] As in the first example, for equivalent thrust, the fan section of engine 1 has a performance coefficient of less than 1.05 (s / m)², while the performance coefficient of engine 2 is equal to 1.11, thus falling between 1.05 and 1.3. Therefore, compared to engine 1, engine 2 can achieve at least an equivalent level of efficiency with less mass, drag, and volume, while being easier to install on an aircraft and maintaining an acceptable level of operability regarding certification.

Claims

1. A fan section (2) of an aircraft propulsion system, said fan section (2) comprising a fan rotor (9) with a performance coefficient greater than or equal to 1.05 and less than or equal to 1.3, wherein, The performance coefficient is defined as follows: Where: Cp is the performance coefficient, with units of seconds per square meter (s / m)²; FPR is the pressure ratio of the fan section (2); Cs is the solidity of the fan rotor (9); and U is the peripheral velocity at the blade tip of the fan rotor (9), in meters per second (m / s). The pressure ratio and the peripheral velocity are measured at cruising speed.

2. The fan section (2) according to claim 1, wherein, The fan pressure ratio is greater than or equal to 1.1 and less than or equal to 1.

45.

3. The fan section (2) according to claim 1 or 2, wherein, The realism is greater than or equal to 0.9 and less than or equal to 1.3, preferably greater than or equal to 0.9 and less than or equal to 1.

1.

4. The fan section (2) according to any one of claims 1 to 3, wherein, The peripheral speed is greater than or equal to 260 m / s and less than or equal to 400 m / s, for example, greater than or equal to 270 m / s and less than or equal to 350 m / s.

5. The fan section (2) according to any one of claims 1 to 4, wherein, The elongation ratio of the fan blade (14) is the ratio of the height of the fan blade to the chord length (c1) at the tip of the fan blade, and the elongation ratio of the fan blade is greater than or equal to 2.4 and less than or equal to 3.

0.

6. The fan section (2) according to any one of claims 1 to 5, wherein, The diameter of the fan rotor (9) is greater than or equal to 177.8 cm and less than or equal to 304.8 cm, for example less than or equal to 304.8 cm, for example about 228.6 cm.

7. The fan section (2) according to any one of claims 1 to 6, wherein, The ratio of the hub to the tip of the fan section is greater than or equal to 0.22 and less than or equal to 0.32, for example greater than or equal to 0.235 and less than or equal to 0.30, or again for example less than or equal to 0.

27.

8. An aircraft propulsion system (1), comprising: - Drive shaft (11), which is rotatable about a rotation axis (X); - Fan shaft (20); - In any one of claims 1 to 7, the fan section (2) is rotatably driven by the fan shaft (20); and - A reduction mechanism (19) connects the drive shaft (11) and the fan shaft (20) so that the fan shaft (20) is driven at a speed lower than that of the drive shaft (11).

9. The aerospace propulsion system (1) according to claim 8 further includes a fan housing surrounding the fan blades (14).

10. The propulsion system (1) according to claim 8 or 9, wherein, The bypass ratio of the propulsion system (1) is greater than or equal to 10, for example, between 10 and 35 and including 10 and 35, for example, between 10 and 18 and including 10 and 18.

11. An aircraft (100) comprising at least one propulsion system (1) according to any one of claims 8 to 10, the propulsion system being attached to the aircraft via a mast.

12. A method for determining the dimensions of a fan section (2) of an aircraft propulsion system, the method comprising the steps of determining the pressure ratio of the fan section, the solidity of the rotor (9) of the fan section, and the peripheral velocity of the rotor (9), such that the performance coefficient of the fan section (2) is greater than or equal to 1.05 and less than or equal to 1.3, wherein, The performance coefficient is defined as follows: Where: Cp is the performance coefficient; FPR is the pressure ratio of the fan section (2); Cs is the solidity of the fan rotor (9); and U is the peripheral velocity at the blade tip of the fan rotor (9); The pressure ratio and the peripheral velocity are measured at cruising speed.

13. The size determination method according to claim 12 further includes the step of determining the size of the fan blade (14) such that the elongation ratio of the fan blade is greater than or equal to 2.4 and less than or equal to 3.0, wherein, The elongation ratio is the ratio of the height of the fan blade (14) to the chord length (c1) at the tip of the fan blade.

14. A method for manufacturing a fan section, comprising the following steps: The size determination method according to claim 12 or 13 is used to determine the size of the fan section; and Manufacture the fan section to such a specific size.

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