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 1.40 * 105 N / m2 and less than or equal to 1.70 * 105 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 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 particularly, to an aircraft propulsion system including a ducted fan and having a high or even very high bypass ratio. Background of the Invention

[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, and 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 in order to obtain aviation components and products with lower energy intensity and a more environmentally friendly nature, the integration and use of which in civil aviation has 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, a propulsion system having 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) has been proposed. In order 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 the 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 the rotational speed 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 increasing 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 accommodated 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 stable 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 aviation propulsion system is proposed, comprising:

[0008] - a drive shaft that is rotatably movable about a rotation 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 is coupled to the drive shaft and the fan shaft 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 1.40×10 5 N / m 2 and less than or equal to 1.70×10 5 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 takeoff rating when the propulsion system is stationary, and is expressed in newtons (N);

[0015] n is the number of blades 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 rotation axis, and is expressed in meters (m).

[0017] Some preferred but non-limiting features of the aviation 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 22.0×10 6 W / m 2 , preferably greater than or equal to 16.0×10 6 W / m 2 and less than or equal to 22.0×10 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 takeoff 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, and the fan compression ratio corresponds to a pressure ratio less than or equal to 1.45 between the outlet of the fan rotor and the inlet of the fan rotor;

[0022] - The diameter of the fan rotor is included between 80 inches and 120 inches (including the end values), for example, about 90 inches;

[0023] - The fan section is ducted, and the bypass ratio of the propulsion system is greater than or equal to 10, for example, included between 10 and 35 (including the end values), preferably, between 10 and 18 (including the end values);

[0024] - The fan section is ducted, and when the propulsion system is stationary at takeoff rating at sea level under standard atmospheric pressure, the circumferential velocity at the tip of the blade of the fan rotor is included between 260 m / s and 400 m / s;

[0025] - The hub-to-tip ratio of the fan rotor is included between 0.22 and 0.32;

[0026] - The propulsion system further includes a driving turbine and a compressor directly connected by a drive shaft, 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 further 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; and / or

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

[0030] 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 pylon.

[0031] 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 ducted fan rotor to drive the fan rotor at a speed lower than the speed of the drive shaft, wherein 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 1.40×10 5 N / m 2 and less than or equal to 1.70×105 N / m 2 , wherein the thrust density of each blade of the fan rotor is defined by the following formula:

[0032]

[0033] 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);

[0034] n is the number of blades in the fan rotor; and

[0035] 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).

[0036] Optionally, 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 22.0×10 6 W / m 2 , preferably greater than or equal to 16.0×10 6 W / m 2 and less than or equal to 22.0×10 6 W / m 2 , wherein the power density of each blade of the fan rotor is defined by the following formula:

[0037]

[0038] where 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).

[0039] According to a fourth aspect, the present application provides a method for manufacturing a propulsion system, including the following steps:

[0040] - determining the size of the propulsion system according to the third aspect; and

[0041] - manufacturing the propulsion system. BRIEF DESCRIPTION OF THE DRAWINGS

[0042] 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 in reference to the accompanying drawings, in which:

[0043] Figure 1 is a schematic partial cross - sectional view of an example of a propulsion system according to an embodiment, in which the fan section is ducted;

[0044] Figure 2 is a schematic cross - sectional view of an example of a star reduction gear;

[0045] Figure 3 is a schematic cross - sectional view of an example of a planetary reduction gear;

[0046] Figure 4 is an example of an aircraft that may include at least one propulsion system according to an embodiment;

[0047] Figure 5 is a flow chart showing an example of steps in a sizing or manufacturing method according to an embodiment of the present invention.

[0048] In all the figures, like elements have the same reference numerals. DETAILED DESCRIPTION

[0049] The propulsion system 1 has a main direction extending along a longitudinal axis X, and when in operation, in the direction of the air flow in the propulsion system 1, from upstream to downstream, it includes: a fan section 2 and a body 3, which is generally 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).

[0050] 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.

[0051] 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 and passing through the axis X. Further, 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 "interior") and "outer" (respectively "exterior") 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.

[0052] 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, and the primary air flow F1 and the secondary air flow F2 circulate from upstream to downstream in the propulsion system 1.

[0053] 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.

[0054] 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 portion 9 of the fan section 2.

[0055] 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 rotation of the rotor of the high-pressure compressor 5 via the high-pressure shaft 10. The rotor of the low-pressure turbine 8 drives the rotation of the rotor of the low-pressure compressor 4 and the rotor portion 9 of the fan section 2 via the low-pressure shaft 11. Thus, 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 the rotational speed 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 rotation of the rotor of the low-pressure compressor 4 via an intermediate shaft. The rotation of the fan rotor 9 and the rotor of the high-pressure compressor 5 are respectively driven by the low-pressure shaft 11 and the high-pressure shaft 10.

[0056] The low-pressure shaft 11 is usually received in a part of its length within 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, the low-pressure shaft 11 and the high-pressure shaft 10 are driven 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, the low-pressure shaft 11 and the high-pressure shaft 10 are driven in opposite directions around the longitudinal axis X. Where appropriate, an intermediate shaft is received 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.

[0057] The fan section 2 at least includes a fan rotor 9 that can be driven by the turbine sections 7, 8 to rotate relative to the stator portion of the propulsion system 1. The fan rotor 9 includes a hub 13 and blades 14 that radially extend from the hub 13. The blades 14 of the 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 are pivotally mounted along the setting axis and are connected to a pitch-changing mechanism 15 installed in the propulsion system 1, and the setting is adjusted by the pitch-changing mechanism 15 according to the flight phase. The pitch-changing mechanism 15 is Figure 1 shown in dashed lines to indicate that this feature is optional.

[0058] 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 having 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 are 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.

[0059] 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 values between 10 and 35 (including the endpoints). It should be noted that when the bypass ratio is greater than or equal to 25, the fan rotor 9 is preferably a variable-setting fan rotor. To calculate the bypass ratio, when the propulsion system 1 is stationary, uninstalled, and at takeoff rating at 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, the 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.

[0060] The fan rotor 9 is decoupled from the low-pressure shaft 11 using a reduction mechanism 19 to facilitate independent optimization of their respective rotational speeds. The reduction mechanism 19 is provided 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 reduction mechanism 19, and the fan shaft 20 connects the outlet of the reduction mechanism 19 to the fan rotor 9. Thus, the fan rotor 9 is driven by the low-pressure shaft 11 via the reduction mechanism 19 and the fan shaft 20 to rotate at a rotational speed lower than that of the low-pressure turbine 8.

[0061] This decoupling makes it possible to reduce the rotational speed and pressure ratio of the fan rotor 9 and increase the power extracted by the low-pressure turbine 8. In fact, the overall efficiency of the propulsion system primarily depends on the propulsion efficiency, which is favorably affected by minimizing the change in 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 consists of 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 related: 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 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 defined position on the inner surface of the inlet flow passage of the fan rotor 9 to the tip 21 of the fan blade 14).

[0062] 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).

[0063] The fan section 2 is ducted and includes a fan casing 12 in which the fan rotor 9 is accommodated. The fan rotor 9 extends upstream of the fan stator 16. The guide vanes 17 of the fan stator 16 are commonly referred to as Outlet Guide Vanes (OGV) and have a fixed arrangement relative to the hub of the fan stator. In addition, the bypass ratio of the propulsion system 1 is preferably greater than or equal to 10, for example, included between 10 and 35 (including the end values), preferably, included between 10 and 18 (including the end values). The circumferential speed at the tip 21 of the blades of the fan rotor 9 can also be included between 260 m / s and 400 m / s. The blades 14 of the fan rotor 9 can be fixed or have a variable setting. Then, the fan pressure ratio can be included between 1.20 and 1.45.

[0064] In the ducted fan section 2, the fan section 2 is not surrounded by a fan casing. 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 of propulsion system 1, the guide vanes 17 of the straightener 16 are fixedly rotated relative to the rotation axis X of the upstream fan rotor 9, and thus are not subject to centrifugal force. The guide vanes 17 of the straightener 16 are also variable-setting guide vanes.

[0065] 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 pinion 19a (the inlet of the reduction mechanism 19), a ring gear 19b (the outlet of the reduction mechanism 19), and a series of planetary gears 19c. The sun pinion 19a is centered on the rotation axis X of the reduction mechanism 19 (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 pinion 19a and is configured to drive the fan shaft 20 to rotate about the rotation axis X. The series of planetary gears 19c are circumferentially distributed around the rotation axis X between the sun pinion 19a and the ring gear 19b, and each planetary gear 19c meshes with the sun pinion 19a internally and with the ring gear 19b externally. The series of planetary gears 19c are mounted on a planetary gear carrier 19d, which is fixed relative to the stator portion 19e of the propulsion system 1, for example, fixed relative to the casings of the compressor sections 4, 5. According to the second variant, the reduction mechanism 19 may 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 planetary gear carrier 19d (the planetary gear carrier can thus rotate relative to the stator portion 19e of the propulsion system 1, for example, rotate relative to the casings of the compressor sections 4, 5).

[0066] Regardless of the structure of the reduction mechanism 19, the diameters of the ring gear 19b and the planet carrier 19d are larger than the diameter of the sun pinion 19a, so that the rotational speed of the fan rotor 9 is lower than the rotational speed of the low-pressure shaft 11.

[0067] The bypass ratio of the reduction mechanism 19 is greater than or equal to 2.5 and less than or equal to 11, preferably greater than or equal to 2.7 and less than or equal to 6.0. Typically, the bypass ratio of the reduction mechanism 19 is about 3.0.

[0068] The redline speed of the low-pressure shaft 11 is included between 8500 rpm and 12000 rpm, preferably between 9000 rpm and 11000 rpm. This redline speed corresponds to the absolute maximum speed that the low-pressure shaft 11 may encounter during the entire flight (in accordance with European Certification Regulation (EASA) CS-E 740 (or in accordance with US Certification Regulation 14-CFR Part 33.87)). This redline 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 redline speed. This redline speed forms part of the data declared in the engine certification (type certificate data sheet). In fact, this rotational speed is typically used as a reference speed for determining the size of the propulsion system 1 and in some certification tests such as blade loss or rotor integrity tests.

[0069] To optimize the performance of the propulsion system 1 in terms of specific fuel 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 1.40×10 5 N / m 2 and less than or equal to 1.70×10 5 N / m 2 , where the thrust density of each blade 14 is defined by the following formula:

[0070]

[0071] where: FN is the thrust generated by the fan rotor 9, expressed in newtons (N);

[0072] n is the number of blades 14 in the fan rotor 9; and

[0073] 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.

[0074] In fact, the applicant has noticed that when the thrust density is less than 1.40×10 5 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 excessive resistance. In addition, when the thrust density is greater than 1.70×10 5 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 gear 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 1.40×10 5 N / m 2 and 1.70×10 5 N / m 2 , thus making it possible to obtain a compromise between the integration and the performance of the propulsion system 1. This thrust density range for 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.

[0075] For example, a propulsion system 1 according to the invention including a ducted fan rotor and having a thrust density of 1.5×10 5 N / m 2 for each of its fan blades has a 5% lower unit consumption compared to the same propulsion system having a thrust density of 2.1×10 5 N / m 2 for each of its fan blades. 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 (at a defined fan diameter), the overall compression ratio and the number of stages in the compression section and the turbine section generally have little or no effect on the thrust density of each blade 14.

[0076] 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), it is possible to size and manufacture the propulsion system 1 to obtain a thrust density of each blade 14 between 1.40×10 5 N / m 2 and 1.70×10 5 N / m 2The power density therebetween. Compared with a propulsion system having a conventional reduction mechanism, the diameter D can be increased, for example, and the pressure ratio of the fan 2 can be decreased. The number (n) of the fan blades 14 and the rotational speed of the fan rotor 9 can also be adjusted to meet the 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 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). 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, etc.): In particular, the flow rate of the gas generator can be decreased and the bypass ratio of the reduction mechanism 19 can be increased.

[0077] For each blade 14 having a thrust density between 1.40×10 5 N / m 2 and 1.70×10 5 N / m 2 the diameter D of the fan rotor can be included between 80 inches (203.2 cm) and 120 inches (304.8 cm) (including the end values), for example, approximately 90 inches (228.6 cm), which enables the integration of the propulsion system 1 in a conventional manner, in particular, under the wing of the aircraft 100.

[0078] In addition, the fan rotor 9 includes at least twelve blades 14 and at most twenty-four blades 14, preferably, at least sixteen blades 14 and at most twenty-two blades 14. The number of the guide vanes 16 in the fan stator 17 depends on the acoustic standard defined for the propulsion system 1 and is at least equal to the number of the blades 14 of the fan rotor 9.

[0079] 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 22.0×10 6 W / m 2 preferably, greater than or equal to 16.0×10 6 W / m 2 and less than or equal to 22.0×10 6 W / m 2 wherein the power density of each blade 14 of the fan rotor 9 is defined by the following formula:

[0080]

[0081] wherein the fan power corresponds to the power of the fan rotor 9 and is expressed in watts (W).

[0082] The fan rotor 9 also has a hub-to-tip ratio that is between 0.22 and 0.32. In the case of a fan rotor with a fixed setting, the hub-to-tip ratio can be between 0.24 and 0.32. In the case of a fan rotor with a variable setting, the hub-to-tip ratio preferably is between 0.26 and 0.32 in order to allow for the integration of the pitch-changing 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 that is the intersection between the leading edge 22 and the surface radially bounding 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 that is the intersection between the leading edge 22 and the tip 21 of the fan blade (and this outer radius R e corresponds 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, a decrease in 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 fan rotor 9 is sized such that its hub-to-tip ratio is between 0.22 and 0.32, in particular such that a thrust density and a power density within the intervals defined above can be obtained for each blade 14.

[0083] A dual-rotor propulsion system 1 having a thrust density and a power density within the intervals defined above for each blade 14 of the fan rotor 9 can in particular 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.

[0084] Comparative example:

[0085] The engine 1 is a dual-rotor propulsion system that includes a ducted fan section 2 corresponding to what needs to be improved in the current state of the art (at the filing date of the present application).

[0086] The engine 2 is a dual-rotor propulsion system 1 that includes a ducted fan section 2 that complies with the teachings of the present application and in which the thrust density of each fan blade is equal to 1.49×10 5 N / m 2 .

[0087]

[0088]

[0089]

[0090] The thrust density of each fan blade of engine 1 is greater than 1.70×10 5 N / m 2 , while for engine 2 the thrust density of each fan blade is included between 1.40×10 5 N / m 2 and 1.70×10 5 N / m 2 . It can be seen that, due to its improved thrust density per blade, engine 2 has a lower specific consumption than engine 1 without the need to increase the inlet temperature of the low-pressure turbine. Thus, the efficiency of the fan section 2 is improved without the need to further load the low-pressure turbine.

[0091] 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 increased, which allows the propulsion efficiency to be increased and a comparable thrust to be maintained for a given reduction in the pressure ratio of fan 2. Additionally, the overall compression ratio is increased and the inlet temperature of the high-pressure turbine 7 is increased, which allows the thermal efficiency of the propulsion system 1 to be increased. Finally, as long as the temperature of the low-pressure turbine 8 remains stable, its mechanical load (N12S) can be increased and thus the number of its stages reduced.

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 1.40×10 5 N / m 2 and less than or equal to 1.70×10 5 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 22.0×10 6 W / m 2 , preferably, greater than or equal to 16.0×10 6 W / m 2 and less than or equal to 22.0×10 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 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 a pressure ratio less than or equal to 1.45 between the outlet and the inlet of the fan rotor (9).

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 120 inches, i.e., 304.8 cm, including 80 inches and 120 inches, for example, about 90 inches, i.e., 228.6 cm.

5. The propulsion system (1) according to any one of claims 1 to 4, wherein, The fan section (2) is ducted, and the bypass ratio of the propulsion system (1) is greater than or equal to 10, for example, the bypass ratio is between 10 and 35, including 10 and 35, preferably, the bypass ratio is between 10 and 18, including 10 and 18.

6. The propulsion system (1) according to any one of claims 1 to 5, wherein, The fan section (2) is ducted, and 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 260 m / s and 400 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.

32.

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. An aircraft (100) comprising at least one propulsion system (1) according to any one of claims 1 to 11, said propulsion system being fixed to the aircraft via a pylon.

13. A method for sizing a propulsion system (1), the propulsion system including a reduction mechanism (19) coupled to a drive shaft (11) and a ducted fan rotor (9) to drive the fan rotor (9) at a speed lower than the speed of the drive shaft (11), wherein, The dimensions of the fan rotor (9) are 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 1.40×10 5 N / m 2 and less than or equal to 1.70×10 5 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 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 axis of rotation, and is expressed in metres m.

14. The dimension determination method according to claim 13, wherein, The dimensions of the fan rotor (9) are 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 22.0×10 6 W / m 2 , preferably greater than or equal to 16.0×10 6 W / m 2 and less than or equal to 22.0×10 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.

15. A method for manufacturing a propulsion system (1) comprising the steps of: - determining the dimensions of the propulsion system (1) according to any one of claims 13 and 14; and - manufacturing the propulsion system (1).

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