Aerial propulsion system with higher propulsion efficiency

Abstract

The invention relates to an aero-propulsion system (1), the thrust density of each blade (14) of the fan rotor (9) of which is greater than or equal to 5.0 * 104 N / m2 and less than or equal to 10.0 * 104 N / m2, where the thrust density of each blade (14) is defined by the following formula: # imgabs0 # where: FN is the thrust generated by the fan rotor (9), and FN is the thrust generated by the fan rotor (9); and FN is measured when the propulsion system (1) is stationary at standard atmospheric pressure and at sea level at a takeoff rating, and is represented by Newton (N); n is the number of blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), D being measured in a plane perpendicular to the rotation axis (X) at the intersection between the tip (21) and the leading edge (22) of the blade (14) of the fan rotor (9), and expressed in meter.

Description

Field of the Invention

[0001] The present application generally relates to the field of propulsion systems and, more specifically, to an aircraft propulsion system including an unducted fan and having a high or even very high bypass ratio. Background Art

[0002] In the direction of the air flow from upstream to downstream, a propulsion system generally includes a fan section, a compressor section, a combustion chamber, and a turbine section. The compressor section may include a low-pressure compressor and a high-pressure compressor. The turbine section may particularly 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 (if applicable) are driven to rotate by the low-pressure turbine via a low-pressure shaft.

[0003] Technical research work has led to a very significant improvement in the environmental performance of aircraft. The applicant has considered factors affecting all design and development stages to obtain aircraft components and products with lower energy intensity and a more environmentally friendly nature, whose integration and use in civil aviation have a moderate environmental impact, so as to achieve the purpose of improving the energy efficiency of the aircraft.

[0004] Therefore, in order to improve the propulsion efficiency of the propulsion system, reduce its specific fuel consumption, and reduce the noise emitted by the fan section, propulsion systems with a high bypass ratio (BPR) (corresponding to the ratio between the flow rate of the secondary air flow and the flow rate of the primary air flow) have been proposed. To achieve such a bypass ratio, the fan section is decoupled from the low-pressure turbine, so that their respective rotational speeds can be optimized independently. Generally, a reduction mechanism placed between the upstream end of the low-pressure shaft and the rotor of the fan section is used to achieve decoupling. Then, the rotor of the fan section is driven by the low-pressure shaft via the reduction mechanism at a rotational speed lower than that of the low-pressure shaft.

[0005] However, increasing the bypass ratio (BPR) and the fan pressure ratio involves increasing the fan diameter and, by extension, increasing the external dimensions of the propulsion system (and thus its mass and drag), which makes the integration of the propulsion system more difficult in addition to increasing its mass and specific fuel consumption. The flow rate in the high-pressure body and the dimensions of the high-pressure body also decrease, which places restrictions on the low-pressure body. In particular, since the low-pressure shaft is housed within the high-pressure shaft, reducing the dimensions of the low-pressure body (and thus reducing the dimensions of the high-pressure shaft) involves reducing the diameters of the high and low pressure shafts, so that the low-pressure shaft becomes supercritical. However, within the operating range of the propulsion system, a supercritical shaft includes bending deformation modes. Then, the dynamics of the low-pressure shaft must be controlled so that these deformation modes do not occur within the stability range, thus avoiding any risk of damaging the propulsion system. Summary of the Invention

[0006] An object of the present application is to optimize the performance of the propulsion system in terms of unit consumption, mass and drag, while ensuring the possibility of integrating the propulsion system into an aircraft.

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

[0008] - a drive shaft that is rotatably movable about a rotational axis;

[0009] - a fan shaft;

[0010] - a fan section that includes a ducted fan rotor driven to rotate by the fan shaft, the fan rotor including a plurality of blades;

[0011] - a reduction mechanism that couples 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 thrust density of each blade of the fan rotor of the propulsion system is greater than or equal to 5.0×10 4 N / m 2 and less than or equal to 10.0×10 4 N / m 2 , wherein the thrust density of each blade is defined by the following formula:

[0013]

[0014] where: FN is the thrust generated by the fan rotor, FN is measured at standard atmospheric pressure and at sea level at take-off rating when the propulsion system is stationary, and is expressed in newtons (N);

[0015] n is the number of blades (14) in the fan rotor; and

[0016] D is the diameter of the fan rotor, D is measured at the intersection between the tip and the leading edge of the blades of the fan rotor in a plane perpendicular to the rotational axis, and is expressed in meters (m).

[0017] Some preferred but non-limiting features of the aircraft propulsion system according to the first aspect are as follows, individually or in combination:

[0018] - The power density of each blade of the fan rotor is greater than or equal to 3.65×10 6 W / m 2 and less than or equal to 7.50×10 6 W / m 2 , wherein the power density of each blade of the fan rotor is defined by the following formula:

[0019]

[0020] Wherein: the power of the fan corresponds to the power of the fan rotor, and the power of the fan is measured at take-off rating at sea level under standard atmospheric pressure when the propulsion system is stationary, and is expressed in watts (W);

[0021] - The fan section also has a fan compression ratio, which corresponds to the pressure ratio between the outlet of the fan rotor and the inlet of the fan rotor, and this pressure ratio is less than or equal to 1.45, preferably less than or equal to 1.30;

[0022] - The diameter of the fan rotor is between 80 inches (203.2 cm) and 185 inches (469.9 cm) (including the end values), preferably between 120 inches (304.8 cm) and 170 inches (431.8 cm) (including the end values), for example, approximately 156 inches (396.2 cm);

[0023] - The bypass ratio of the propulsion system is greater than or equal to 40, for example, between 40 and 80 (including the end values);

[0024] - When the propulsion system is stationary at take-off rating at sea level under standard atmospheric pressure, the circumferential velocity at the tip of the blades of the fan rotor is between 210 m / s and 260 m / s;

[0025] - The hub-to-tip ratio of the fan rotor is between 0.22 and 0.34;

[0026] - The propulsion system also includes a driving turbine and a compressor directly connected by a drive shaft, and the driving turbine includes at least three stages and at most five stages;

[0027] - The compressor includes at least two stages and at most four stages;

[0028] - The propulsion system also includes a high-pressure turbine and a high-pressure compressor connected via a high-pressure shaft, the high-pressure shaft rotates faster than the drive shaft, and the high-pressure turbine is a two-stage turbine;

[0029] - The high-pressure compressor includes at least eight stages and at most eleven stages; and / or

[0030] The fan rotor includes at least ten fan blades and at most eighteen fan blades, preferably at least twelve fan blades and at most sixteen fan blades.

[0031] According to a second aspect, the present application proposes an aircraft including at least one propulsion system according to the first aspect, and the propulsion system is fixed to the aircraft via a mast.

[0032] According to a third aspect, the present application provides a method for determining the size of a propulsion system, the propulsion system including a reduction mechanism coupled to a drive shaft and a fan rotor to drive the ducted fan rotor at a speed lower than the speed of the drive shaft. The size of the fan rotor is determined such that the thrust density of each blade of the fan rotor of the propulsion system is greater than or equal to 5.0×10 4 N / m 2 and less than or equal to 10.0×10 4 N / m 2 , wherein the thrust density of each blade of the fan rotor is defined by the following formula:

[0033]

[0034] where: FN is the thrust of the fan rotor, and FN is measured when the propulsion system is stationary at takeoff rating under standard atmospheric pressure and at sea level, and is expressed in Newtons (N);

[0035] n is the number of blades in the fan rotor; and

[0036] D is the fan diameter, D is measured at the intersection between the tip and the leading edge of the blade of the fan rotor in a plane perpendicular to the axis of rotation, and is expressed in meters (m).

[0037] Optionally, the size of the fan rotor (9) can be further determined such that the power density of each blade of the fan rotor is greater than or equal to 3.65×10 6 W / m 2 and less than or equal to 7.50×10 6 W / m 2 , wherein the power density of each blade of the fan rotor is defined by the following formula:

[0038]

[0039] wherein the power of the fan rotor is measured when the propulsion system is stationary at takeoff rating under standard atmospheric pressure and at sea level, and is expressed in watts (W).

[0040] According to a fourth aspect, a method for manufacturing a propulsion system is provided, including the following steps:

[0041] - determining the size of the propulsion system according to the third aspect; and

[0042] - manufacturing the propulsion system. Description of the Drawings

[0043] Other features, objects, and advantages of the present invention will become more apparent from the following description, which is illustrative and non - limiting and should be read with reference to the accompanying drawings, in which:

[0044] Figure 1 is a schematic partial cross - sectional view of an example of a propulsion system according to an embodiment, where the fan section is non - ducted;

[0045] Figure 2 is a schematic cross - sectional view of an example of a star reduction gear;

[0046] Figure 3 is a schematic cross - sectional view of an example of a planetary reduction gear;

[0047] Figure 4 is an example of an aircraft that may include at least one propulsion system according to an embodiment;

[0048] Figure 5 is a flowchart showing an example of steps in a sizing or manufacturing method according to an embodiment.

[0049] In all the figures, like elements have the same reference numerals. Detailed Description

[0050] The propulsion system 1 has a main direction extending along the longitudinal axis X, and when it operates, in the direction of the airflow in the propulsion system 1 from upstream to downstream, it includes: a fan section 2 and a body 3, the body 3 ( ) is commonly referred to as a "gas generator", and the body 3 includes a compressor section 4, 5, a combustion chamber 6, and a turbine section 7, 8. Here, the propulsion system 1 is an aviation propulsion system 1, which is configured to be fixed to the aircraft 100 via a pylon (or mast).

[0051] The compressor sections 4, 5 include a series of stages, each stage including a bladed wheel (rotor) 4a, 5a rotating in front of a guide vane (stator) 4b, 5b. The turbine sections 7, 8 also include a series of stages, each stage including a guide vane (stator) 7b, 8b, and a bladed wheel (rotor) 7a, 8a rotating after the guide vane.

[0052] In the present application, the axial direction corresponds to the direction of the longitudinal axis X, the direction of the longitudinal axis X corresponds to the rotation of the gas generator shaft, and the radial direction is the direction perpendicular to this axis X and passing through this axis X. In addition, the circumferential (or transverse or tangential) direction corresponds to the direction perpendicular to the longitudinal axis X and not passing through the longitudinal axis X. Unless otherwise specified, the terms "inner" (respectively, internal) and "outer" (respectively, external) are used relative to the radial direction, such that the inner or inner surface of an element is closer to the axis X than the outer or outer surface of the same element.

[0053] 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 from upstream to downstream in the propulsion system 1.

[0054] The secondary air flow F2 (also referred to as the "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.

[0055] The primary air flow F1 flows in the main passage within the body 3, successively passing through the compressor sections 4, 5, the combustion chamber 6, and the turbine sections 7, 8. In the combustion chamber 6, the primary air flow F1 is mixed with fuel to be used as an oxidant. The primary air flow F1 causes the rotation of the rotors of the turbine sections 7, 8 by receiving energy from the combustion chamber 6, which in turn drives the rotation of the rotors of the compressor sections 4, 5 and the rotor part 9 of the fan section 2.

[0056] In the twin-rotor propulsion system 1, the compressor sections 4, 5 may include a low-pressure compressor 4 and a high-pressure compressor 5. The turbine sections 7, 8 may include a high-pressure turbine 7 and a low-pressure turbine 8. The rotor of the high-pressure turbine 7 drives the rotor of the high-pressure compressor 5 to rotate via the high-pressure shaft 10. The rotor of the low-pressure turbine 8 drives the rotor of the low-pressure compressor 4 and the rotor part 9 (propeller) of the fan section 2 to rotate via the low-pressure shaft 11. Accordingly, the body 3 includes a high-pressure body and a low-pressure body. The high-pressure body includes the high-pressure compressor 5, the high-pressure turbine 7, and the high-pressure shaft 10. The low-pressure body 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 body is greater than that of the low-pressure body. In the triple-rotor propulsion system 1, the turbine sections 7, 8 further include an intermediate turbine located between the high-pressure turbine 7 and the low-pressure turbine 8, 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 kept driven respectively by the low-pressure shaft 11 and the high-pressure shaft 10.

[0057] The low-pressure shaft 11 is generally accommodated in a part of the length of the high-pressure shaft 10 and is coaxial with the high-pressure shaft 10. The low-pressure shaft 11 and the high-pressure shaft 10 can rotate together, that is, drive the low-pressure shaft 11 and the high-pressure shaft 10 in the same direction around the longitudinal axis X. As a variant, the low-pressure shaft 11 and the high-pressure shaft rotate in opposite directions, that is, drive the low-pressure shaft 11 and the high-pressure shaft 10 in opposite directions around the longitudinal axis X. Where appropriate, the intermediate shaft is accommodated 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.

[0058] The fan section 2 at least includes 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 includes a hub 13 and blades 14 radially extending from the hub 13. The blades 14 of each rotor 9 can be fixed relative to the hub 12 or can have a variable setting. In this case, the roots of the blades 14 of each rotor 9 are pivotally mounted along the setting axis and connected to a pitch-changing mechanism 15 installed in the propulsion system 1, and the setting is adjusted by this pitch-changing mechanism 15 according to the flight phase. The pitch-changing mechanism 15 is shown in Figure 1 in dashed lines to indicate that this feature is optional.

[0059] The fan section 2 can also include a fan stator 16 or a straightener which includes guide vanes 17 mounted on the hub 18 of the fan stator 16 and has the function of straightening the secondary air flow F2 flowing at the outlet of the fan rotor 9. The guide vanes 17 of the fan stator 18 can be fixed relative to the hub 18 or can have a variable setting. In a manner similar to the rotor blades 14, the roots of the stator guide vanes 17 are pivotally mounted along the setting axis X and connected to a pitch-changing mechanism 15a which is generally different from the pitch-changing mechanism of the fan rotor 9, and the setting is adjusted by the pitch-changing mechanism according to the flight phase.

[0060] In order to improve the propulsion efficiency of the propulsion system 1 and reduce its specific fuel consumption as well as the noise emitted by the fan section 2, the propulsion system 1 has a high bypass ratio. The high bypass ratio mentioned here means that the bypass ratio is greater than or equal to 10, for example, including between 10 and 80 (including the endpoints). In order to calculate the bypass ratio, when the propulsion system 1 is stationary, uninstalled and at takeoff rating under standard atmospheric pressure (as defined by the International Civil Aviation Organization (ICAO) Manual, Doc 7488 / 3, 3rd Edition) and at sea level (this condition is called Seal Level Standard (SLS)), the mass flow rate of the secondary air flow F2 and the mass flow rate of the primary air flow F1 are measured. It should be noted that in this application, parameters (pressure, flow rate, thrust, speed, etc.) are systematically determined under these conditions. The "uninstalled" here means the measurement is carried out when the propulsion system 1 is on a test bench (and not installed on the aircraft 100), so that the measurement is easier to perform.

[0061] The deceleration mechanism 19 is used to decouple the fan rotor 9 from the low-pressure shaft 11 so as to independently optimize their respective rotational speeds. The deceleration mechanism 19 is arranged between the upstream end of the low-pressure shaft 11 and the fan rotor 9. In this case, the propulsion system 1 further includes an additional shaft called the fan shaft 20. The low-pressure shaft 11 connects the low-pressure turbine 8 to the inlet of the deceleration mechanism 19, and the fan shaft 20 connects the outlet of the deceleration mechanism 19 to the fan rotor 9. Therefore, the fan rotor 9 is driven by the low-pressure shaft 11 via the deceleration mechanism 19 and the fan shaft 20 to rotate at a rotational speed lower than that of the low-pressure turbine 8.

[0062] This decoupling enables the rotational speed and pressure ratio of the fan rotor 9 to be reduced and the power extracted by the low-pressure turbine 8 to be increased. In fact, the overall efficiency of the propulsion system primarily depends on the propulsion efficiency, which is favorably affected by minimizing the change in the kinetic energy of the air flowing through the propulsion system 1. In a propulsion system 1 with a high bypass ratio, most of the flow that generates the propulsion force is constituted by the secondary air flow F2 of the propulsion system 1, and the kinetic energy of the secondary air flow F2 is mainly affected 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 pressure ratio of the fan (which corresponds to the ratio between the average pressure at the outlet of the fan stator 17 (or, in the absence of a stator, the fan rotor 9) and the average pressure at the inlet of the fan rotor 9) is less than or equal to 1.70, preferably less than or equal to 1.50, for example, included between 1.05 and 1.45. Here, the average pressure is measured along the height of the blade 14 (from the radially defining position of the inner surface of the inlet passage of the fan rotor 9 to the tip 21 of the fan blade 14).

[0063] The propulsion system 1 is configured to provide a thrust included between 18000 lbf (80068 N) and 51000 lbf (222411 N), preferably, the thrust is included between 20000 lbf (88964 N) and 35000 lbf (15568 N).

[0064] The fan section 2 can be non-ducted. In the case of a ducted fan section 2, the fan section 2 includes a fan housing 12, and the fan rotor 9 is accommodated in the fan housing 12.

[0065] In the ducted fan section 2, the fan section 2 is not surrounded by a fan housing. Since the fan section 2 is ducted, the blades 14 of the fan rotor 9 have a variable setting. A propulsion system including at least one ducted fan rotor 9 is also referred to as an "open rotor" or "ducted fan". The propulsion system 1 may include two ducted and counter-rotating fan rotors 9. Such a propulsion system 1 is referred to as a "Contra-Rotating Open Rotor (CROR)" or "Unducted Double Fan (UDF)". One or more fan rotors 9 may be arranged at the rear of the main body 3 to be a pusher type, or arranged at the front of the main body 3 to be a tractor type. As a variant, the propulsion system 1 may include a single ducted fan rotor 9 and a ducted fan stator 16 (straightener). Such a propulsion system 1 is referred to as an "Unducted Single Fan (USF)". In the case of the USF type propulsion system 1, the guide vanes 17 of the straightener 16 rotate fixedly relative to the rotation axis X of the upstream fan rotor 9, so it is not subject to centrifugal force. The guide vanes 17 of the straightener 16 are also guide vanes with variable settings.

[0066] There is no fairing around the fan section 2, which allows a very significant increase in the bypass ratio without affecting the propulsion system 1 by the mass of the housing or engine nacelle designed to surround the fan section 2. Therefore, the bypass ratio of the propulsion system 1 including the non-ducted fan section 2 is greater than or equal to 40, for example, between 40 and 80 (including the endpoints). The circumferential speed at the tip 21 of the blades 14 of one or more fan rotors 9 can also be between 210 m / s and 260 m / s. Then, the fan pressure ratio can preferably be between 1.05 and 1.20.

[0067] The reduction mechanism 19 may include, for example, a reduction mechanism 19 having a planetary gear train, such as a single-stage or two-stage reduction mechanism of the "planetary" or "star" type. According to the first variant, the reduction mechanism 19 may be of the star type ( Figure 2) and includes a sun gear 19a (the inlet of the reduction mechanism 19), a ring gear 19b (the outlet of the reduction mechanism 19), and a series of planet gears 19c. The sun gear 19a is centered on the rotation axis X of the reduction mechanism 19 (substantially coinciding with the longitudinal axis X) and is configured to be driven to rotate by the low-pressure shaft 11. The ring gear 19b is coaxial with the sun gear 19a and is configured to drive the fan shaft 20 to rotate about the rotation axis X. The series of planet gears 19c are circumferentially distributed around the rotation axis X between the sun gear 19a and the ring gear 19b, and each planet gear 19c meshes with the sun gear 19a internally and with the ring gear 19b externally. The series of planet gears 19c are mounted on a planet gear carrier 19d, and the planet gear carrier 19d is fixed relative to the stator portion 19e of the propulsion system 1, for example, fixed to the housing of the compressor sections 4, 5. According to the second variant, the reduction mechanism 19 can be planetary ( Figure 3 ), in which case the ring gear 19b is fixedly mounted on the stator portion 19e of the propulsion system 1, and the fan shaft 20 is driven to rotate by the planet gear carrier 19d (the planet gear carrier can thus rotate relative to the stator portion 19e of the propulsion system 1, for example, rotate relative to the housing of the compressor sections 4, 5).

[0068] Regardless of the structure of the reduction mechanism 19, the diameters of the ring gear 19b and the planet gear carrier 19d are larger than the diameter of the sun gear 19a, so that the rotational speed of the fan rotor 9 is lower than the rotational speed of the low-pressure shaft 11.

[0069] The reduction ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11, preferably between 9.0 and 11.0.

[0070] The red line speed of the low-pressure shaft 11 is included between 8500 rpm and 12000 rpm, preferably between 9000 rpm and 11000 rpm. This red line speed corresponds to the absolute maximum speed that the low-pressure shaft 11 may encounter during the entire flight (according to the European Certification Regulation (EASA) CS-E 740 (or according to the US Certification Regulation 14-CFR Part 33.87)). This red line speed corresponds to the maximum rotational speed when the propulsion system is intact (and possibly at the end of its life). Therefore, under flight conditions, the low-pressure shaft 11 is likely to reach this red line speed. This red line speed constitutes part of the data declared in the engine certification (type certificate data sheet). In fact, this rotational speed is usually used as a reference speed for determining the size of the propulsion system 1 and some certification tests (such as blade loss or rotor integrity tests).

[0071] In order to optimize the performance of the propulsion system 1 in terms of unit consumption, mass, and drag, while ensuring the possibility of integrating the propulsion system 1 into the aircraft 100, the thrust of each blade 14 of the fan rotor 9 is greater than or equal to 5.0×10 4 N / m 2 and less than or equal to 10.0×10 4 N / m 2 , where the thrust density of each blade 14 is defined by the following formula:

[0072]

[0073] where: FN is the thrust generated by the fan rotor 9, expressed in newtons (N);

[0074] n is the number of blades 14 in the fan rotor 9; and

[0075] D is the diameter of the fan rotor 9, and is expressed in meters (m). This diameter 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. It should be noted that since Figure 1 is a partial view, the diameter D is only partially visible.

[0076] When the propulsion system 1 includes two fan rotors 9, the thrust density of each blade 14 of the fan rotor is less than or equal to 4.0×10 4 N / m 2 .

[0077] In fact, the applicant has noticed that when the thrust density is less than 5.0×10 4 N / m 2 , it is difficult to integrate the propulsion system 1 because it is too large in volume, too heavy in mass, and generates too much drag. In addition, when the thrust density is greater than 10.0×10 4 N / m 2 , the performance of the propulsion system 1 in terms of unit consumption decreases. Therefore, when the propulsion system 1 includes a reduction mechanism 19 and has a high bypass ratio, the size of the propulsion system 1 is such that the thrust density of each blade 14 of the fan rotor 9 is included between 5.0×10 4 N / m 2 and 10.0×10 4 N / m 2 , thus making it possible to obtain a compromise between the integration and the performance of the propulsion system 1. This interval of thrust density of each blade 14 is also compatible with a fan pressure ratio of less than 1.45, which makes it possible to optimize the propulsion efficiency of the propulsion system 1.

[0078] For example, a non-ducted fan rotor according to the invention, and the thrust density of each fan blade 14 thereof is equal to 6.5×104 N / m 2 The propulsion system 1, whose thrust density per each fan blade is equal to 4.5×10 4 N / m 2 compared to the same propulsion system, has a fan diameter that is 13% smaller, which helps integrate the propulsion system 1 under the wing without affecting the efficiency of the propulsion system 1. The thrust density of each blade of the propulsion system 1 is affected first-order by the diameter D of the fan rotor and the pressure ratio of the fan section 2. The bypass ratio, the overall compression ratio, and the number of stages in the compression and turbine sections generally have little or no effect on the thrust density of each blade 14.

[0079] Therefore, by first fixing the thrust (FN) to be generated by the fan section 2 and modifying the diameter (D) of the fan rotor (thereby modifying the pressure ratio of the fan section 2) to obtain such thrust, the sizing and manufacture of the propulsion system 1 can be achieved to obtain a thrust density for each blade 14 that is included between 5.0×10 4 N / m 2 and 10.0×10 4 N / m 2 Compared to a propulsion system with a conventional reduction mechanism, the diameter D can be slightly reduced, for example, to allow the integration of the propulsion system under the wing, and the pressure ratio of the fan 2 can be adjusted accordingly to obtain the desired thrust. The number (n) of fan blades 14 and the rotational speed of the fan rotor 9 can also be adjusted to meet performance, acoustic, and integration requirements. Depending on the aerodynamic characteristics of the fan section 2, the propulsion system 1 can be modified to integrate variable pitch mechanisms 15, 15a, enabling the adjustment of the settings of the blades 14 of the rotor 9 of the fan section 2 (and possibly the guide vanes 16 of the stator 17). Additionally, depending on the integration performance balance (the fuel consumption balance of the propulsion system 1 integrated into the aircraft (mass, unit consumption, drag)) and aircraft constraints (in terms of integration and program constraints), the fan section 2 can include a single fan rotor 9 or two counter-rotating fan rotors 9. Finally, the thermodynamic cycle is adjusted to adapt to the various parameters of the propulsion system 1 and thus determine the size of the propulsion system 1 (fan diameter, number of blades, pressure ratio of the fan section 2, overall compression ratio, etc.): In particular, the flow rate of the gas generator can be reduced and the bypass ratio of the reduction mechanism 19 can be increased.

[0080] For a value included between 5.0×10 4 N / m 2 and 10.0×10 4 N / m 2For the thrust density of each blade 14 therebetween, the diameter D of the fan rotor may be included between 80 inches (203.2 cm) and 185 inches (469.9 cm) (including the end values), preferably greater than or equal to 100 inches (254 cm), for example, between 120 inches (304.8 cm) and 170 inches (431.8 cm), for example, 156 inches (396.2 cm), which allows the integration of the propulsion system 1 in a conventional manner, in particular the integration of the propulsion system 1 under the wing of an aircraft.

[0081] Furthermore, the fan rotor 9 includes at least ten blades 14, at most eighteen blades 14, preferably at least twelve blades 14, at most sixteen blades 14. The number of guide vanes 16 in the fan stator 17 depends on the acoustic standard defined for the propulsion system 1 and is less than the number of blades 14 of the fan rotor 9.

[0082] To further improve the propulsion efficiency of the propulsion system 1, the power density of each blade 14 of the fan rotor 9 is greater than or equal to 3.65×10 6 W / m 2 and less than or equal to 7.50×10 6 W / m 2 , where the power density of each blade 14 of the fan rotor 9 is defined by the following formula:

[0083]

[0084] where the fan power corresponds to the power of the fan rotor 9 and is expressed in watts (W).

[0085] The fan rotor 9 also has a hub-to-tip ratio included between 0.22 and 0.34, which allows the integration of the variable pitch mechanism 15. The hub-to-tip ratio corresponds to the ratio between the inner radius R i of the fan rotor 9 and the outer radius R e . The inner radius R i corresponds to the distance between the axis of rotation X and an intersection point, which is the intersection point between the leading edge 22 and the surface radially delimited within the flow passage at the inlet of the fan rotor 9 (and this inner radius R i corresponds to the connection point of the leading edge 22 with the aerodynamic surface of the platform of the fan rotor 9). The outer radius R e corresponds to the distance between the axis of rotation X and an intersection point, which is the intersection point between the leading edge 22 and the tip 21 of the fan blade (and this outer radius R e(corresponding to half of the fan diameter D). The lower the hub-to-tip ratio of the fan rotor 9, the higher the efficiency of the fan rotor 9. However, the reduction of the hub-to-tip ratio of the fan rotor 9 means an increase in the mechanical load on the hub 13 of the fan rotor 9. The size of the fan rotor 9 is determined such that its hub-to-tip ratio is included between 0.22 and 0.34, and in particular, such that a thrust density and a power density of each blade 14 within the intervals defined above can be obtained.

[0086] The dual-rotor propulsion system 1 having a thrust density and a power density of each blade 14 of the fan rotor 9 within the intervals defined above may particularly include a two-stage high-pressure turbine 7, a high-pressure compressor 5 including at least eight and at most eleven stages, a low-pressure turbine 8 including at least three and at most five stages, and a low-pressure compressor 4 including at least two and at most four stages.

[0087] Comparative example:

[0088] The engine 1 is a dual-rotor propulsion system corresponding to the current state of the art to be improved (at the filing date of the present application), which includes a non-ducted fan section 2.

[0089] The engine 2 is a dual-rotor propulsion system 1 in accordance with the teachings of the present application, which includes a non-ducted fan section, two shafts, and the thrust density of each fan blade is equal to 6.53×10 4 N / m 2 .

[0090]

[0091]

[0092] The thrust density of each fan blade of the engine 1 is greater than 4.7×10 4 N / m 2 , while for the engine 2 the thrust density of each fan blade is included between 5.0 N / m 2 and 10.0×10 4 N / m 2 . It can be seen that, due to its improved thrust density of each fan blade, the engine 2 is more compact and has a smaller mass compared to the engine 1. In this comparative example, the reduction of the mass of the fan section 2 is estimated to be about 20%. Considering that the fan section represents one third of the mass of the propulsion system 1, this corresponds to a reduction of about 7% of the mass of the propulsion system 1 and has an impact on the aircraft (cantilever mass, reduction of the diameter of the fan rotor 9, etc.). Thus, with the inlet temperature of the low-pressure turbine 8 and the fan thrust being equal, the engine 2 can be more easily installed on the aircraft 100.

[0093] To move from (reference) engine 1 to engine 2 (in line with the present disclosure), the fan diameter D and the bypass ratio BPR are reduced, which allows for an increase in the integration level of engine 2. However, the pressure ratio in the fan section of engine 2 increases slightly (but remains below 1.45) to maintain equal 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 main 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, its mechanical load (N12 S) can be increased to reduce the number of its stages.

Claims

1. An aircraft propulsion system (1), comprising: - A drive shaft (11) that is rotatably movable about a rotational axis (X); - A fan shaft (20); - A fan section (2) that includes a ducted fan rotor (9) driven to rotate by the fan shaft (20), the fan rotor (9) including a plurality of blades (14); - A reduction mechanism (19) that couples the drive shaft (11) and the fan shaft (20) to drive the fan shaft (20) at a rotational speed lower than that of the drive shaft (11); The thrust density of each blade (14) of the fan rotor (9) of the propulsion system (1) is greater than or equal to 5.0×10 4 N / m 2 and less than or equal to 10.0×10 4 N / m 2 , where the thrust density of each blade (14) is defined by the following formula: Wherein: FN is the thrust generated by the fan rotor (9), FN is measured when the propulsion system (1) is stationary at takeoff rating under standard atmospheric pressure and at sea level, and is expressed in newtons N; n is the number of blades (14) in the fan rotor (9); and D is the diameter of the fan rotor (9), D is measured at the intersection between the tip (21) and the leading edge (22) of the blades (14) of the fan rotor (9) in a plane perpendicular to the rotational axis (X), and is expressed in meters m.

2. The propulsion system (1) according to claim 1, wherein, The power density of each blade (14) of the fan rotor (9) is greater than or equal to 3.65×10 6 W / m 2 and less than or equal to 7.50×10 6 W / m 2 , where the power density of each blade (14) of the fan rotor (9) is defined by the following formula: Wherein, the power of the fan corresponds to the power of the fan rotor (9), the power of the fan is measured when the propulsion system (1) is stationary at takeoff rating under standard atmospheric pressure and at sea level, and is expressed in watts W.

3. The propulsion system (1) according to any one of claims 1 and 2, wherein, The fan section (2) also has a fan compression ratio that corresponds to the pressure ratio between the outlet and the inlet of the fan rotor (9), the pressure ratio being less than or equal to 1.45, preferably less than or equal to 1.

30.

4. The propulsion system (1) according to any one of claims 1 to 3, wherein The diameter of the fan rotor (9) is between 80 inches, i.e., 203.2 cm, and 185 inches, i.e., 469.9 cm, inclusive of 80 inches and 185 inches, preferably between 120 inches, i.e., 304.8 cm, and 170 inches, inclusive of 120 inches and 170 inches, for example, approximately 156 inches, i.e., 396.2 cm.

5. The propulsion system (1) according to any one of claims 1 to 4, wherein, The bypass ratio of the propulsion system (1) is greater than or equal to 40, for example, the bypass ratio is between 40 and 80, inclusive of 40 and 80.

6. The propulsion system (1) according to any one of claims 1 to 5, wherein, When the propulsion system (1) is stationary at takeoff rating under standard atmospheric pressure and at sea level, the circumferential speed at the tip (21) of the blades (14) of the fan rotor (9) is between 210 m / s and 260 m / s.

7. The propulsion system (1) according to any one of claims 1 to 6, wherein, The hub-to-tip ratio of the fan rotor (9) is between 0.22 and 0.

34.

8. The propulsion system (1) according to any one of claims 1 to 7, further comprising: A drive turbine (8) and a compressor (4) directly connected by the drive shaft (11), the drive turbine (8) including at least three stages and at most five stages.

9. The propulsion system (1) according to claim 8, wherein, The compressor (4) includes at least two stages and at most four stages.

10. The propulsion system (1) according to any one of claims 1 to 9, further comprising: A high-pressure turbine (7) and a high-pressure compressor (5) connected via a high-pressure shaft (10), the high-pressure shaft (10) rotating faster than the drive shaft (11), the high-pressure turbine (7) being a two-stage turbine.

11. The propulsion system (1) according to claim 10, wherein, The high-pressure compressor (5) includes at least eight stages and at most eleven stages.

12. The propulsion system (1) according to any one of claims 1 to 11, wherein, The fan rotor (9) includes at least ten fan blades (14), at most eighteen fan blades (14), preferably at least twelve fan blades (14), at most sixteen fan blades (14).

13. An aircraft (100) includes at least one propulsion system (1) according to any one of claims 1 to 12, the propulsion system being fixed to the aircraft via a pylon.

14. A method for sizing a propulsion system (1), the propulsion system including a reduction gear (19) coupled to a drive shaft (11) and a fan rotor (9) to drive the non-ducted fan rotor (9) at a speed lower than the speed of the drive shaft (11), wherein, The size of the fan rotor (9) is determined such that the thrust density of each blade (14) of the fan rotor (9) of the propulsion system (1) is greater than or equal to 5.0×10 4 N / m 2 and less than or equal to 10.0×10 4 N / m 2 , where the thrust density of each blade (14) of the fan rotor (9) is defined by the following formula: where: FN is the thrust of the fan rotor (9), and FN is measured when the propulsion system (1) is stationary at take-off rating at standard atmospheric pressure and at sea level, and is expressed in newtons N; n is the number of blades (14) in the fan rotor (9); and D is the fan diameter, D being measured in a plane perpendicular to the axis of rotation at the intersection between the tip (21) and the leading edge (22) of the blades (14) of the fan rotor (9), and is expressed in metres m.

15. The dimension determination method according to claim 14, wherein, The size of the fan rotor (9) is further determined such that the power density of each blade (14) of the fan rotor (9) is greater than or equal to 3.65×10 6 W / m 2 and less than or equal to 7.50×10 6 W / m 2 , wherein the power density of each blade (14) of the fan rotor (9) is defined by the following formula: wherein the power of the fan rotor (9) is measured when the propulsion system (1) is stationary at take-off rating at standard atmospheric pressure and at sea level, and is expressed in watts W.

16. A method for manufacturing a propulsion system (1) includes the following steps: - determining the dimensions of the propulsion system (1) according to any one of claims 14 and 15; and - manufacturing the propulsion system (1).

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