Aerial propulsion system with improved propulsion efficiency

Abstract

The invention relates to an aviation propulsion system (1) comprising:-a fan shaft (20); a fan section (2) comprising a fan rotor (22) rotated by a fan shaft; -a main body (3) comprising a compressor section (29) comprising a low-pressure compressor (4) and a high-pressure compressor (5), a combustion chamber (6) and a turbine section (30) comprising a high-pressure turbine (7) rotating the high-pressure compressor (5) by means of a high-pressure shaft (10) and a low-pressure turbine (8) rotating the low-pressure compressor (4) by means of a low-pressure shaft (11); -a speed reduction mechanism (19) coupling the low-pressure shaft (8) and the fan shaft (20) so as to drive the fan shaft (20) at a speed lower than the rotational speed of the low-pressure shaft (8); the low-pressure compressor (4) comprises a first stage (44) and at least one other stage (46, 48), each stage sequentially comprising a rotor (44r, 46r, 48r) and a stator (44s, 46s, 48s) in the direction of flow of the gas, the value of the maximum speed of the low-pressure compressor (4) being in the range of 8500 revolutions per minute (rpm) to 10000 revolutions per minute (rpm), the value of the ratio between the average radius (Rmoy) of the low-pressure compressor and the inner radius (Ric) of the rotor of the first stage being in the range of 1.1 to 1.4.

Description

Technical Field

[0001] This disclosure generally relates to the field of propulsion systems, and more specifically to aerospace propulsion systems including ducted or unducted fans and having high or even very high bypass ratios. Background Technology

[0002] An aerospace propulsion system typically includes a fan section, a compressor section, a combustion chamber, and a turbine section in the direction of gas flow from upstream to downstream. The compressor section may include a low-pressure compressor and a high-pressure compressor, and the turbine section may include a high-pressure turbine and a low-pressure turbine.

[0003] When the propulsion system is running, the high-pressure compressor is driven to rotate by a high-pressure turbine via a high-pressure shaft. The fan and an optional low-pressure compressor are driven to rotate by a low-pressure turbine via a low-pressure shaft.

[0004] One of the goals of technological research is to improve the environmental performance of aircraft. This is why relevant factors are taken into account at every stage of design and development to obtain components and aviation products with lower energy intensity and a more environmentally conscious approach. The integration of these components and aviation products and their use in civil aviation have modest environmental consequences, with the ultimate goal of improving the energy efficiency of aircraft.

[0005] Therefore, to improve the propulsion efficiency of the propulsion system and reduce its specific energy consumption and noise emitted by the fan section, propulsion systems with a high bypass ratio (BPR) have been proposed. This BPR corresponds to the ratio between the secondary airflow rate and the primary airflow rate. To achieve this bypass ratio, the fan section can be decoupled from the low-pressure turbine, thereby independently optimizing their respective rotational speeds. Typically, this decoupling is achieved through a reduction gear located between the upstream end of the low-pressure shaft and the rotor of the fan section. Thus, the rotor of the fan section is driven by the low-pressure shaft at a lower speed than that of the low-pressure shaft via the reduction gear.

[0006] However, in such a system, the overall performance of the low-pressure compressor may be limited by dimensional constraints, particularly those resulting from the space required to install bearings in the chamber between the low-pressure shaft and the moving wheels of the low-pressure compressor. The volume occupied in this chamber can depend particularly on the selected rotational speed of the compressor and the power selected to be obtained via the compressor, combustion chamber, and turbine. The volume of the reduction gear also affects the dimensional constraints that influence the efficiency of the low-pressure compressor located within the flow channel and upstream of the chamber. Furthermore, considering the volume of the flow channel in the low-pressure compressor, the dimensional constraints will depend on the available space for integrating the compressor within its environment, the flow channel volume being contained between the inner and outer walls of the flow channel. Therefore, the performance of the low-pressure compressor depends on aerodynamics and the space occupied around the compressor. Summary of the Invention

[0007] One objective of this invention is to optimize the performance of the propulsion system, particularly to improve the compactness of the low-pressure compressor while maintaining a high bypass ratio.

[0008] Therefore, in a first aspect, an aviation propulsion system is proposed, which includes: - Fan shaft; - Fan section, the fan section includes a fan rotor that rotates by a fan shaft; - Main body, the main body includes compressor section, combustion chamber and turbine section. The compressor section includes low-pressure compressor and high-pressure compressor. The turbine section includes high-pressure turbine and low-pressure turbine. The high-pressure turbine rotates the high-pressure compressor through high-pressure shaft. The low-pressure turbine rotates the low-pressure compressor through low-pressure shaft. -Reduction mechanism, which connects the low-pressure shaft and the fan shaft, thereby driving the fan shaft at a speed lower than that of the low-pressure shaft; The low-pressure compressor includes a first stage and at least one other stage, each stage comprising a rotor and a stator in the direction of gas flow. The maximum speed of the low-pressure compressor is in the range of 8,500 rpm to 10,000 rpm, and the ratio between the average radius Rmoy of the low-pressure compressor and the inner radius of the first-stage rotor is in the range of 1.1 to 1.4.

[0009] Therefore, the low-pressure compressor offers optimized performance, particularly optimized efficiency. Specifically, compared to existing propulsion system low-pressure compressors, this compressor exhibits lower aerodynamic load and lower speed. This allows for better integration of the reduction gear and low-pressure compressor into the propulsion system without degrading the overall engine performance.

[0010] The compressor's average radius (Rmoy) is defined as the average of the radii of each rotor, corresponding to the distance between the propulsion system's rotation axis X and a point selected at the midpoint of the height and distance between the leading and trailing edges of the rotor. The inner radius Ric corresponds to the radius of the first rotor (i.e., the upstream rotor), and is the distance between the propulsion system's rotation axis and a point selected at the blade root at the leading edge of the first-stage rotor.

[0011] Specifically, in order to achieve optimal efficiency for the low-pressure compressor (or BP compressor), its load or aerodynamic load can preferably be less than or equal to 0.33, and the compressor load is defined by the following equation:

[0012] in: H is the difference in air enthalpy between the inlet and outlet of the BP compressor, expressed in joules per kilogram (J / kg); N1 is the number of stages in the BP compressor; and Umoy is the average tangential velocity at 50% of the rotor height of the BP compressor, expressed in meters per second (m / s). Umoy can be calculated using the following equation:

[0013] Where Rmoy corresponds to the average radius of the rotor at 50% height, and Rcomp is the compressor speed in revolutions per minute (tr / min).

[0014] The aerodynamic load (or load) of a compressor corresponds to the aerodynamic load of the compressor's rotor stage and to the ratio between the working demand (compressed air) and the compressor's operating characteristics / conditions. Under equivalent geometry and the same operating conditions, a higher load value results in greater compression, but also lower efficiency and greater difficulty in operating the compressor. This is why the compressor load value is preferably no more than 0.33.

[0015] The difference in enthalpy between the outlet and inlet air of a low-pressure compressor corresponds to the difference in enthalpy between the leading edge of the first rotor and the trailing edge of the rectifier (or stator) located downstream of the last rotor.

[0016] The tangential velocity of the airflow (Umoy) is defined as the average tangential velocity of each stage of the low-pressure compressor at 50% height.

[0017] In addition, the Mach number at the base of the leading edge of the rotor of the first stage of the BP compressor can preferably be 1 ± 10% to avoid transonic shock waves.

[0018] Some preferred, but non-limiting, features of the propulsion system according to the first aspect are as follows, which are employed individually or in combination: - The diameter (D) of the fan rotor is in the range of 1.270 m to 4.300 m; this diameter is measured at the leading edge of each blade of the fan rotor; with this diameter, the propulsion system can be integrated into a conventional aircraft without modifying the aircraft, especially without modifying the size of the landing gear or wings, or strengthening the wings. - The fan section is ductless, and the diameter (D) of the fan rotor is in the range of 1.980m to 4.300m; this diameter is measured at the leading edge of each blade of the propeller; - The fan section is ducted, and the diameter of the fan rotor is in the range of 1.270m to 3.048m; this diameter is measured at the leading edge of each blade of the propeller; the diameter can be, for example, 2.286m; - The low-pressure compressor is a three-stage compressor, and the compression ratio of the low-pressure compressor is in the range of 2.8 to 4.8. The compression ratio is defined as the ratio between the gas pressure leaving the low-pressure compressor and the gas pressure entering the low-pressure compressor; the compression ratio of the BP compressor allows the compression ratio of each stage to be in the range of 1.4 to 1.7, so that the compressor can operate with optimized efficiency. Each stage compression ratio is defined as the compression ratio of the compressor at one power (1 / n); the compression ratio is specifically measured using a sensor comb in the test engine, specifically by measuring flow rate and pressure under climb conditions up to 35,000 feet, at Mach 0.8 and ISA+15; - The low-pressure compressor is a two-stage compressor, and the compression ratio of the low-pressure compressor is in the range of 1.7 to 2.5. The compression ratio is defined as the ratio between the gas pressure leaving the low-pressure compressor and the gas pressure entering the low-pressure compressor; the compression ratio of the BP compressor enables the compression ratio of each stage to be in the range of 1.4 to 1.7, so that the compressor can operate with optimized efficiency. - The bypass ratio of the propulsion system is in the range of 12 to 16. The bypass ratio is defined as the ratio between the mass flow rate of the secondary airflow (F2) flowing through the fan section around the main body and the mass flow rate of the main airflow (F1) flowing through the main body; this value is preferably applicable to ducted propellers. - The inner radius of the first-stage rotor is in the range of 250 mm to 350 mm; this range of radius values ​​is particularly effective for engines with short to medium thrust (SMR) (especially from 100,000 to 175,000 Newtons), and these values ​​are measured when the propulsion system is stationary at takeoff rating in standard atmosphere and at sea level. - The reduction ratio of the reduction mechanism is in the range of 2.5 to 11, preferably in the range of 2.7 to 6.0, and more preferably in the range of 2.7 to 3.6; for example, the reduction ratio can be 3.0; - When the propulsion system is stationary at sea level in standard atmosphere at takeoff ratings, the outer circumferential velocity at the tip of the fan rotor blades is in the range of 270 m / s to 380 m / s; the fan section is preferably ducted. The hub-to-tip ratio of the fan rotor is in the range of 0.22 to 0.32, preferably in the range of 0.235 to 0.300, for example, about 0.27. The hub-to-tip ratio is defined as the ratio between the inner radius (Ri) and the outer radius (Re) of the fan rotor. - The ratio between the value of the first chord measured at the tip of the fan rotor blade and the value of the second chord measured at the root of the fan rotor blade is in the range of 0.56 to 0.80; - The consistency of the fan rotor at the tip of the fan rotor blades is in the range of 1.0 to 1.3. The consistency is defined as the ratio between the value of the first chord measured at the tip of the blade and the value of the inter-blade spacing. - The fan rotor has multiple blades, the number of which is in the range of 16 to 24, for example 22; in particular, the number of fan blades is optimized by taking into account the fan diameter, the density at the tip of the fan rotor blades and the value of the peripheral velocity at the tip of the fan rotor blades as described above. - The fan section includes a fan stator having a plurality of outlet guide vanes, the number of which is in the range of 36 to 52, preferably in the range of 38 to 44, for example 40; in particular, the number of fan stator guide vanes is optimized taking into account the fan diameter, the density at the tip of the fan rotor blades and the value of the peripheral velocity at the tip of the fan rotor blades (especially the values ​​described above); - The low-pressure turbine comprises multiple stages, with the number of stages ranging from 3 to 5; this number of stages ensures optimized aerodynamic load of the low-pressure turbine and makes it possible to prevent excessive turbine mass and / or volume. - The high-pressure turbine consists of two stages; - High-pressure compressors consist of multiple stages, with the number of stages ranging from 8 to 11. - The power density of the low-pressure compressor (4) is in the range of 0.64 to 0.97, and the power density of the compressor is defined by the following equation:

[0019] Wherein: BPR is the bypass ratio of the propulsion system, and BPR is measured when the propulsion system is stationary at takeoff rating in standard atmosphere and at sea level; N et Um is the number of stages in the compressor; Um is the average peripheral velocity of the compressor rotor measured in standard atmosphere and at sea level at takeoff rating, expressed in pounds of thrust and meters per second (m / s); Rext_moy is the average outer radius of the compressor, which is equal to the arithmetic mean of the outer radii of the low-pressure compressor rotor. The outer radius corresponds to the distance between the rotor tip and the axis of rotation at the midpoint between the leading and trailing edges of the rotor, expressed in meters (m).

[0020] In summary, the number of stages described for high-pressure turbines and low-pressure turbines, as well as high-pressure compressors and low-pressure compressors, provides a sufficient number of stages for the turbines and compressors to perform their functions, but the number of stages cannot be too high so that the turbines or compressors are not too heavy or too large.

[0021] For the same reason, blade size and the number of blades are also important parameters.

[0022] In the second aspect, an aircraft is also proposed, which includes a propulsion system according to the first aspect.

[0023] In a third aspect, a design method or manufacturing method for a propulsion system according to the first aspect is also proposed, the method comprising the following steps: - For low-pressure compressors, select a maximum speed value within the range of 8500 rpm to 10000 rpm, and - Select the average radius (Rmoy) and inner radius (Ric) of the low-pressure compressor such that the ratio between the average radius (Rmoy) of the low-pressure compressor and the inner radius (Ric) of the first-stage rotor is in the range of 1.1 to 1.4.

[0024] This design approach is also applicable to propulsion systems that include the preferred features described above.

[0025] This design method may specifically include the following steps, in which, for the low-pressure compressor (4), the power density value is selected to be in the range of 0.64 to 0.97, and the power density of the compressor is defined by the following equation:

[0026] Wherein: BPR is the bypass ratio of the propulsion system, and BPR is measured when the propulsion system is stationary at takeoff rating in standard atmosphere and at sea level; N et Um is the number of stages in the compressor; Um is the average peripheral velocity of the compressor rotor measured when the propulsion system is stationary at sea level and in standard atmosphere at takeoff rating, in meters per second (m / s); Rext_moy is the average outer radius of the compressor, which is equal to the arithmetic mean of the outer radii of the low-pressure compressor rotor. The outer radius corresponds to the distance between the rotor tip and the axis of rotation at the midpoint between the leading and trailing edges of the rotor, in meters (m). Attached Figure Description

[0027] Other features and advantages of the invention will become apparent from the following description, given only by way of example and with reference to the accompanying drawings, in which: - Figure 1 The diagram schematically illustrates an aircraft including a propulsion system. - Figure 2 A partial cross-sectional view of an example of a propulsion system in which the fan is a ducted type is schematically given; - Figure 3 A partial cross-sectional view of an example of a propulsion system in which the fan is ductless is schematically given; - Figure 4 A first example of a planetary speed reduction mechanism is schematically shown; - Figure 5 A first example of a planetary deceleration mechanism is schematically shown; - Figure 6 A partial perspective top view of the fan rotor is schematically given; and - Figure 7 A partial cross-sectional view of an example of a low-pressure compressor is schematically shown. Detailed Implementation

[0028] exist Figure 1 In the example shown, the aircraft is an airplane 100 comprising a fuselage 101 and two wings 102. In this example, the aircraft includes two propulsion systems 1, each propulsion system 1 being connected via a pylon to a corresponding wing 102 of the aircraft 100. In another embodiment, the aircraft may include one or more propulsion systems connected to the fuselage 101.

[0029] Figure 2 A partial cross-sectional view of a first example of propulsion system 1 is schematically shown.

[0030] In this example, propulsion system 1 is a dual-body gas turbine engine with a ducted fan.

[0031] exist Figure 2 In this system, the propulsion system 1 has a main direction extending along the longitudinal axis X. The propulsion system 1 includes a fan section 2 and a main body 3, which is commonly referred to as a "gas generator".

[0032] Fan section 2 includes a fan 22 and a fan housing 12. Fan 22 includes a fan rotor 9. Fan housing 12 surrounds fan rotor 9. Fan rotor 9 is rotatably mounted relative to fan housing 12.

[0033] The fan rotor 9 includes a fan hub 13 and fan blades 14 extending radially from the hub 13. The fan blades 14 may be fixed relative to the fan hub 13 or may have a variable pitch. In the latter case, each fan blade of the fan blades 14 is pivotally mounted on the fan hub 13 along a pitch axis and connected to a pitch-changing mechanism installed in the propulsion system 1. The pitch-changing mechanism allows adjustment of the pitch angle of the fan blades 14 according to the flight phase.

[0034] The fan rotor 9 includes at least fourteen fan blades 14 and no more than twenty-four fan blades 14, preferably at least sixteen fan blades 14 and no more than twenty-four fan blades 14, for example, twenty-two fan blades.

[0035] Furthermore, in this example, fan section 2 also includes a fan stator 16 fixedly mounted to the fan housing 12. The fan stator 16 includes multiple outlet guide vanes (OGVs) 17. The function of this vane assembly is to rectify and regulate the airflow flowing downstream of the fan rotor 9 to improve engine thrust. This vane assembly also acts as a noise reducer.

[0036] Alternatively, the outlet guide vane 17 may have a variable pitch. Optionally, and similar to the fan blades 14 of the fan rotor 9, the root of the outlet vane 17 is pivotally mounted along the pitch axis and connected to a pitch changing mechanism (not shown), the pitch of which is adjusted according to the flight phase.

[0037] The number of outlet guide vanes 17 depends on the acoustic criteria defined for the propulsion system and is at least equal to the number of fan blades 14. The number of outlet guide vanes 17 can range from 36 to 52, preferably from 38 to 44, and for example, 40.

[0038] Figure 3 A second example of propulsion system 1 is schematically shown in a partial cross-sectional view.

[0039] exist Figure 3 In, with Figure 2 Components of the propulsion system that are identical or similar are indicated by the same reference numerals.

[0040] exist Figure 3 In the example shown, propulsion system 1 is a dual-body gas turbine engine with a ductless fan. This dual-body gas turbine engine can be an open rotor type or an unducted single fan type gas turbine engine.

[0041] and Figure 2 Unlike the first example, the fan rotor 9 (also referred to as the "propeller") is not surrounded by the fan housing.

[0042] Fan section 2 is ductless, and fan blade 14 has variable pitch.

[0043] Alternatively, the propulsion system 1 may include two ductless counter-rotating fan rotors 9. The propulsion system 1 is referred to as a contra-rotating open rotor (CROR) turbine or an unducted double fan (UDF) turbine. The fan rotors 9 may be arranged at the rear of the main body 3 as a thruster type, or at the front of the main body 3 as a puller type.

[0044] Without ducts surrounding the fan rotor 9, the bypass ratio can be significantly increased without adversely affecting the propulsion system 1 due to the mass of the casing 12 or nacelle intended to surround the fan section 2. Therefore, the propulsion system 1 including the ductless fan section 1 has a bypass ratio greater than or equal to 40, for example, between 40 and 80 (inclusive).

[0045] The rest of the instruction manual involves Figure 2 The first example of a propulsion system is shown, but the second example of a propulsion system also applies.

[0046] The main body 3 includes a compressor section 29, a combustion chamber 6, and a turbine section 30.

[0047] Compressor section 29 includes low-pressure compressor 4 and high-pressure compressor 5.

[0048] The low-pressure compressor 4 includes a rotor 41 that can be driven to rotate relative to the housing 31 of the propulsion system 1 and a stator 42 that is fixedly mounted on the housing 31.

[0049] The rotor 41 of the low-pressure compressor 4 includes a movable wheel 4a, and the stator 42 of the low-pressure compressor 4 includes a fixed wheel 4b. The movable wheel 4a and the fixed wheel 4b alternate to form a series of low-pressure compressor stages.

[0050] Similarly, the high-pressure compressor 5 includes a rotor 51 that can be driven to rotate relative to the housing 31 of the propulsion system 1 and a stator 52 fixedly mounted on the housing 31.

[0051] The rotor 51 of the high-pressure compressor 5 includes a movable wheel 5a, and the stator 52 of the high-pressure compressor 5 includes a fixed wheel 5b. The movable wheel 5a and the fixed wheel 5b alternate to form a series of high-pressure compressor stages.

[0052] Turbine section 30 includes high-pressure turbine 7 and low-pressure turbine 8.

[0053] The high-pressure turbine 7 includes a rotor 71 that can be driven to rotate relative to the housing 31 of the propulsion system 1 and a stator 72 that is fixedly mounted on the housing 31.

[0054] The rotor 71 of the high-pressure turbine 7 includes a movable wheel 7a, and the stator 72 of the high-pressure turbine 7 includes a fixed wheel 7b. The movable wheel 7a and the fixed wheel 7b alternate to form a series of high-pressure turbine stages.

[0055] Similarly, the low-pressure turbine 8 includes a rotor 81 that can be driven to rotate relative to the housing 31 of the propulsion system 1 and a stator 82 fixedly mounted on the housing 31.

[0056] The rotor 81 of the low-pressure turbine 8 includes a movable wheel 8a, and the stator 82 of the low-pressure turbine 8 includes a fixed wheel 8b. The movable wheel 8a and the fixed wheel 8b alternate to form a series of low-pressure turbine stages.

[0057] The propulsion system 1 includes a low-pressure shaft 11 that connects the rotor 41 of the low-pressure turbine 4 to the rotor 81 of the low-pressure compressor 8. The low-pressure shaft 11 is rotatably mounted relative to the housing 31 about the longitudinal axis X.

[0058] When the propulsion system 1 is working, the rotor 81 of the low-pressure turbine 8 drives the rotor 41 of the low-pressure compressor to rotate via the low-pressure shaft 11.

[0059] The propulsion system 1 further includes a fan shaft 20 and a reduction gear 19. The fan rotor 9 is fixedly mounted on the fan shaft 20. The reduction gear 19 has an input section and an output section. The input section of the reduction gear 19 is connected to the low-pressure shaft 11, and the output section of the reduction gear 19 is connected to the fan shaft 20. Therefore, when the propulsion system 1 is running, the rotor 81 of the low-pressure turbine 8 drives not only the rotor 41 of the low-pressure compressor 4 to rotate, but also the fan rotor 9 to rotate, via the low-pressure shaft 11, the reduction gear 19, and the fan shaft 20.

[0060] The deceleration mechanism 19 may include a single-stage or two-stage planetary deceleration mechanism or a planetary deceleration mechanism.

[0061] For example, Figure 4 The planetary reduction mechanism 19 is shown. The reduction mechanism 19 includes a sun gear 19a (the input of the reduction mechanism 19), a ring gear 19b (the output of the reduction mechanism 19), and a series of planetary gears 19c. The sun gear is centered on the rotational axis of the reduction mechanism 19 (approximately coinciding with the longitudinal axis X) and is configured to be driven to rotate by the low-pressure shaft 11. The ring gear is coaxial with the sun gear 19a and is configured to drive the fan shaft 20 to rotate about its rotational axis X. A series of planetary gears are circumferentially distributed between the sun gear 19a and the ring gear 19b around the rotational axis X of the rotor 9 of the fan section 2. Each planetary gear 19c meshes internally with the sun gear 19a and externally with the ring gear 19b. 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, relative to the housing of the compressor sections 4 and 5.

[0062] In another example, Figure 5 A planetary reduction gear 19 is shown, wherein a 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 support frame 19d.

[0063] Regardless of the construction of the reduction mechanism 19, the diameters of the ring gear 19b and the planetary support frame 19d are larger than the diameter of the sun gear 19a, which means that the rotational speed of the rotor 9 in the fan section 2 is lower than the rotational speed of the low-pressure shaft 11.

[0064] The reduction ratio of the reduction mechanism 19 ranges from 2.5 to 11, preferably from 2.7 to 6.0, more preferably from 2.7 to 3.6, and for example has a value of about 3.0.

[0065] Therefore, the reduction gear 19 provides independent control over the speed of the fan 22 as well as the speed of the low-pressure turbine 8 and the low-pressure compressor 4.

[0066] The low-pressure turbine 8, low-pressure shaft 11, low-pressure compressor 4, fan shaft 20, reduction gear 19 and fan 22 together form the low-pressure body of propulsion system 1.

[0067] The propulsion system 1 further includes a high-pressure shaft 10, which connects the rotor 51 of the high-pressure turbine 5 to the rotor 71 of the high-pressure compressor 7. The high-pressure shaft 10 is rotatably mounted relative to the housing 31 about a longitudinal axis X. The high-pressure shaft 10 is coaxial with and extends about a low-pressure shaft 11.

[0068] When the propulsion system 1 is running, the rotor 71 of the high-pressure turbine 7 drives the rotor 51 of the low-pressure compressor 5 to rotate via the low-pressure shaft 11.

[0069] The high-pressure turbine 7, the high-pressure shaft 10, and the high-pressure compressor 4 together form the high-pressure body of the propulsion system 1.

[0070] The low-pressure shaft 11 and the high-pressure shaft 10 can rotate in the same direction, that is, they are driven around the longitudinal axis X in the same direction of rotation. As a variation, the low-pressure shaft 11 and the high-pressure shaft 10 can rotate in opposite directions, that is, they are driven around the longitudinal axis X in opposite directions of rotation.

[0071] The dual-body propulsion system 1 may specifically include a single-stage high-pressure turbine 7 (i.e., exactly one stage, such as...). Figure 3 (as shown in the example) or a two-stage high-pressure turbine 7 (i.e., exactly including two stages (such as...) Figure 2 (as shown in the example). Preferably, the high-pressure turbine 7 is two-stage.

[0072] High-pressure compressor 5 includes at least eight stages (e.g. Figure 2 (as shown in the example) and no more than eleven levels.

[0073] The low-pressure turbine 8 includes at least three stages (such as...) Figure 2 (as shown in the example) and no more than seven stages. Preferably, the low-pressure turbine 8 includes at least three stages and no more than five stages.

[0074] The low-pressure compressor 4 includes at least two stages and no more than four stages. Preferably, the low-pressure compressor 4 includes two or three stages.

[0075] When the propulsion system is running, the airflow F entering the propulsion system 1 passes through the fan 22 and is then divided into the main airflow F1 and the secondary airflow F2, which flow from upstream to downstream within the propulsion system 1.

[0076] The secondary airflow F2, also known as the bypass airflow, flows around the main body 3 in the secondary flow channel. The secondary airflow F2 enables the cooling of the outer periphery of the main body 3 and is used to generate most of the thrust provided by the propulsion system 1.

[0077] The main airflow F1 flows within the main body 3 in the main flow channel 29, successively passing through the compressor section 29 (low-pressure compressor 4 and high-pressure compressor 5), the combustion chamber 6, and the turbine section 30 (high-pressure turbine 7 and low-pressure turbine 8). In the combustion chamber, the main airflow mixes with the fuel to act as an oxidant. The main airflow F1 passes through the turbine section 30, which receives energy from the combustion chamber 6, causing the movable wheels 7a and 8a of the turbine section 30 to rotate, which in turn drives the movable wheels 4a and 5a of the compressor section 29 and the fan rotor 9 to rotate.

[0078] To improve the propulsion efficiency of propulsion system 1 and reduce the specific energy consumption of the propulsion system and the noise emitted by fan section 2, propulsion system 1 has a high bypass ratio. "High" bypass ratio refers to a bypass ratio greater than or equal to 10 (e.g., in the range of 10 to 80, preferably in the range of 10 to 35, and more preferably in the range of 12 to 16). The bypass ratio is defined as the ratio between the mass flow rate of the secondary airflow F2 and the mass flow rate of the primary airflow F1, measured at takeoff ratings and at sea level in a standard atmosphere (e.g., as defined in the International Civil Aviation Organization (OACI) manual (document 7488 / 3, 3rd edition)). "Not installed" means that the measurement is performed when propulsion system 1 is on a test bench (and not installed in the aircraft), thus making the measurement easier to perform.

[0079] It should be noted that in this application, some parameters are determined under cruise conditions, i.e., at an altitude of 10,668 m (35,000 feet), at Mach 0.8, and under International Standard Atmosphere (ISA, defined by standard ISO 2533 / 1975 version 1975 / Annex 1985). Furthermore, when propulsion system 1 is cooled, i.e., after propulsion system 1 has been shut down for a sufficient period to allow the propulsion system components to cool to ambient temperature, distances (length, radius, diameter, etc.) are measured at ambient temperature (approximately 20°C). In this case, these dimensions change very little compared to the conditions under which propulsion system 1 is at its takeoff rating.

[0080] Including the deceleration mechanism 19 (e.g.) Figure 2 In the propulsion system of the reduction gear shown, the decoupling between the rotational speed of fan 22 and the rotational speed of low-pressure turbine 8 allows for a reduction in the rotational speed and pressure ratio of fan rotor 9, which increases the power obtained by low-pressure turbine 8. The overall efficiency of propulsion system 1 is primarily controlled by propulsion efficiency, which is advantageously influenced by minimizing the change in kinetic energy of air as it passes through propulsion system 1. In propulsion systems with high bypass ratios, the majority of the flow rate generating propulsion is constituted by the secondary airflow F2 of propulsion system 1, the kinetic energy of which is essentially affected by the compression experienced by the secondary airflow F2 as it passes through fan section 2. Therefore, the pressure ratio of fan section 2 and propulsion efficiency are related: the lower the pressure ratio of fan section 2, the better the propulsion efficiency. To improve the propulsion efficiency of propulsion system 1, the fan pressure ratio (corresponding to the ratio between the average pressure output from fan stator 16 (or, if there is no stator 16, from fan rotor 9) and the average pressure input into fan rotor 9) must be less than or equal to 1.70, preferably less than or equal to 1.50, for example, between 1.05 and 1.30. The average pressure is measured at the height of at least one fan blade in the fan blade 14 (i.e., radially and internally defined on the surface of the airflow entering the fan rotor 9 at the tip 21 of the fan blade 14).

[0081] When propulsion system 1 is stationary, not installed, in standard atmosphere (as defined by the International Civil Aviation Organization (OACI) manual, document 7488 / 3, 3rd edition) at takeoff rating and at sea level, propulsion system 1 is configured to provide thrust between 18,000 lbf (80068 N) and 51,000 lbf (222411 N) (e.g., between 20,000 lbf (88964 N) and 35,000 lbf (155688 N)).

[0082] When the propulsion system is stationary at sea level in standard atmosphere at takeoff ratings, the peripheral velocity at the tip of the fan rotor blades ranges from 270 m / s to 380 m / s.

[0083] The diameter D of the fan rotor can range from 1.270 m to 4.300 m. This diameter allows the propulsion system to be conventionally integrated into the aircraft, particularly under the wings of aircraft 1. If the fan section is ductless, the diameter D is preferably in the range of 1.980 m to 4.300 m. If the fan section is ducted, the diameter is preferably in the range of 1.270 m to 3.048 m.

[0084] The fan rotor 22 also has a hub-to-tip ratio in the range of 0.22 to 0.32 (preferably in the range of 0.235 to 0.300, for example, about 0.27). The hub-to-tip ratio corresponds to the ratio between the inner radius Ri and the outer radius Re of the fan rotor 22. The inner radius Ri corresponds to the distance between the axis of rotation X and the point where the leading edge 14' intersects with the surface that radially and internally defines the airflow entering the fan rotor 22 (and corresponds to the point where the leading edge 14' engages with the aerodynamic surface of the platform of the fan rotor 22). The outer radius Re corresponds to the distance between the axis of rotation X and the point where the leading edge 40 intersects with the tip 42 of the fan blade 14 (and corresponds to half the fan diameter D). The lower the hub-to-tip ratio, the better the performance of the fan rotor 22. However, a decrease in the hub-to-tip ratio of the fan rotor 22 leads to an increase in the mechanical load on the hub 13 of the fan rotor 22.

[0085] The ratio between the first chord C1 measured at the tip of the fan rotor blade and the second chord C2 measured at the root of the fan rotor blade is in the range of 0.56 to 0.80. Figure 6 As shown, the chord corresponds to a straight line running from the leading edge toward the trailing edge in a direction perpendicular to the radial axis of the blade.

[0086] Furthermore, the consistency value of the fan rotor 22 at the blade tip 42 is in the range of 1.0 to 1.3. The consistency is defined as the ratio between the value of the first chord C1 measured at the blade tip and the value of the inter-blade spacing P. The inter-blade spacing P corresponds to the distance between two blades measured at the blade root, such as the distance between the leading edges of two consecutive blades.

[0087] To optimize the performance of the propulsion system, especially the efficiency of the low-pressure compressor, the maximum speed of the low-pressure compressor is in the range of 8,500 rpm to 10,000 rpm, and the ratio between the average radius of the low-pressure compressor and the inner radius of the first-stage rotor is in the range of 1.1 to 1.4. The first stage of the low-pressure compressor is the upstream stage of the compressor.

[0088] exist Figure 7 In one embodiment shown, the low-pressure compressor includes a first stage 44 having a rotor 44r and a stator 44s, a second stage 46 having a rotor 46r and a stator 46s, and a third stage 48 having a rotor 48r and a stator 48s.

[0089] The average radius Rmoy of the compressor is defined as the radius R of each rotor. 44 R 46 and R 48The average value of the radius corresponds to the distance between the rotation axis X of the propulsion system and a point selected at the midpoint of the height and distance between the leading and trailing edges of the rotor blade. The inner radius Ric corresponds to the radius of the first rotor 44r, which is the distance between the rotation axis X and a point selected at the root of blade 44r at the midpoint of the distance between the leading and trailing edges of the rotor. The value of the inner radius Ric is in the range of 250 mm to 350 mm.

[0090] Specifically, in order to obtain optimized efficiency of the low-pressure compressor 4, the thermodynamic load or load of the low-pressure compressor is preferably less than or equal to 0.33, and the compressor load is defined by the following equation:

[0091] in: H is the difference in air enthalpy between the outlet and inlet of the low-pressure (BP) compressor, expressed in joules per kilogram (J / kg); N1 is the number of stages in the BP compressor; and Umoy is the average tangential velocity at 50% of the rotor height of the BP compressor, expressed in meters per second (m / s). Umoy can be calculated using the following equation:

[0092] Here, Rmoy corresponds to the previously defined average radius, and Rcomp is the compressor speed in revolutions per minute (tr / min).

[0093] In addition, the Mach number at the base of the leading edge of the rotor of the first stage of the BP compressor is preferably 1 ± 10% to prevent transonic shock waves.

[0094] When the low-pressure compressor 4 is as follows Figure 6 When the three-stage compressor is shown, the compression ratio of each stage is in the range of 2.8 to 4.8. The compression ratio of the BP compressor allows for a compression ratio per stage in the range of 1.4 to 1.7. In another embodiment, not shown, the low-pressure compressor 4 includes two stages, with a compression ratio of each stage in the range of 1.7 to 2.5. The compression ratio of the BP compressor allows for a compression ratio per stage in the range of 1.4 to 1.7.

[0095] Therefore, regardless of the number of stages, the compressor operates with optimized efficiency.

[0096] Alternatively, the power density of the low-pressure compressor 4 is in the range of 0.64 to 0.97, and the power density of the compressor is defined by the following equation:

[0097] Wherein: BPR is the bypass ratio of the propulsion system, measured when the propulsion system is stationary at takeoff rating in standard atmosphere at sea level; it is expressed in pounds of thrust. et Um is the number of stages in the compressor; Um is the average peripheral velocity of the compressor rotor measured when the propulsion system is stationary at takeoff rating in standard atmosphere and at sea level, in pounds of thrust and meters per second (m / s); Rext_moy is the average outer radius of the compressor, which is equal to the arithmetic mean of the outer radii of the low-pressure compressor rotor, corresponding to the distance between the rotor tip and the axis of rotation at the midpoint between the leading and trailing edges of the rotor, and is in meters (m).

[0098] The present invention also includes a method for designing the above-described propulsion system, the propulsion system particularly comprising: - Fan shaft; - Fan section, the fan section includes a fan rotor that is driven to rotate by a fan shaft; - Main body, the main body includes a compressor section, a combustion chamber and a turbine section. The compressor section includes a low-pressure compressor and a high-pressure compressor. The turbine section includes a high-pressure turbine and a low-pressure turbine. The high-pressure turbine drives the high-pressure compressor to rotate through a high-pressure shaft. The low-pressure turbine drives the low-pressure compressor to rotate through a low-pressure shaft. -Reduction mechanism, which connects the low-pressure shaft and the fan shaft, thereby driving the fan shaft at a speed lower than that of the low-pressure shaft; The low-pressure compressor includes a first stage and at least one other stage, each stage comprising a rotor and a stator in the direction of gas flow. The method includes the following steps: - For low-pressure compressors, select a maximum speed value within the range of 8500 rpm to 10000 rpm; and - Select the average radius (Rmoy) and inner radius (Ric) of the low-pressure compressor such that the ratio between the average radius (Rmoy) of the low-pressure compressor and the inner radius (Ric) of the first-stage rotor is in the range of 1.1 to 1.4.

[0099] This design approach is applicable to propulsion systems that include the additional features described earlier, which can be used individually or in combination.

[0100] The design method may specifically include the following steps, in which, for the low-pressure compressor 4, the power density value is selected to be in the range of 0.64 to 0.97, and the power density of the compressor is defined by the following equation:

[0101] Wherein: BPR is the bypass ratio of the propulsion system, and BPR is measured when the propulsion system is stationary at takeoff rating in standard atmosphere and at sea level; N et Um is the number of stages in the compressor; Um is the average peripheral velocity of the compressor rotor measured when the propulsion system is stationary at sea level in standard atmosphere and at takeoff rating, in meters per second (m / s); Rext_moy is the average outer radius of the compressor, which is equal to the arithmetic mean of the outer radii of the low-pressure compressor rotor, corresponding to the distance between the tip of the rotor and the axis of rotation at the midpoint between the leading and trailing edges of the rotor, in meters (m).

[0102] Example The advantages of the present invention are illustrated in a comparative example of two propulsion systems having the characteristics given in the table below.

[0103]

[0104] The first engine (engine 1) is a reference engine that does not possess all the characteristics of the propulsion system of the present invention. The second engine (engine 2) possesses the characteristics of the propulsion system of the present invention. Engine 1 and engine 2 are dual-body gas turbine engines including reduction gears.

[0105] In the table above: -T / O stands for take-off, therefore T / O thrust corresponds to the thrust during the take-off phase; -BPR stands for bypass ratio, which is the bypass ratio of a propulsion system; -OPR stands for Overall pressure ratio, which corresponds to the overall compression ratio of the compressor; -GR represents the gear ratio, corresponding to the reduction ratio; -BP indicates low pressure (basse pression); -HP indicates high pressure (haute pressure); -CBP stands for low-pressure compressor (compresseur basse pression). -CHP stands for high-pressure compressor (compresseur haute pression); -TBP stands for low-pressure turbine. -THP stands for Turbine Haute Pressure. -N2S corresponds to the following product: (velocity of the body of the stage under consideration) 2 × surface area of ​​the stage under consideration; -RHP1 corresponds to the first stage of the high-pressure turbine; -RHP2 corresponds to the second stage of the high-pressure turbine; -RBP1 corresponds to the first stage of the low-pressure turbine; -RBP corresponds to the stage of the low-pressure turbine; -EGT represents the exhaust gas temperature; and -T41 is the inlet temperature of the HP turbine wheel.

[0106] In engine 1, the low-pressure compressor operates at high speeds and high aerodynamic loads to optimize overall dimensions. Therefore, the inner radius of the low-pressure compressor is reduced. Furthermore, the gear reducer is relatively bulky, increasing the engine length. Additionally, the flow channels have small inner radii, making the integration of the low-pressure compressor and determining the location of the bearings supporting it more difficult.

[0107] Advantageously, engine 2 includes a low-pressure compressor with lower aerodynamic load and lower speed. This allows for better integration of the gear reducer and the low-pressure compressor without degrading the overall performance of the engine. Furthermore, the mass loss of the low-pressure turbine with its larger radius is offset by the ease of integration of the low-pressure compressor and the Mach reduction at the head of the low-pressure turbine, thus advantageously affecting performance.

Claims

1. An aeronautical propulsion system (1) comprising: - a fan shaft (20); - a fan section (2) comprising a fan rotor (22) rotated by the fan shaft; - a main body (3) comprising a compressor section (29) comprising a low pressure compressor (4) and a high pressure compressor (5), a combustion chamber (6) and a turbine section (30) comprising a high pressure turbine (7) rotating the high pressure compressor (5) through a high pressure shaft (10) and a low pressure turbine (8) rotating the low pressure compressor (4) through a low pressure shaft (11); - a reduction mechanism (19) coupling the low pressure shaft (8) and the fan shaft (20) so as to drive the fan shaft (20) at a lower speed than the rotation speed of the low pressure shaft (8); the low pressure compressor (4) comprising a first stage (44) and at least one other stage (46, 48), each stage comprising in succession in the direction of the gas flow a rotor (44r, 46r, 48r) and a stator (44s, 46s, 48s), the maximum speed of the low pressure compressor (4) having a value in the range 8500 revolutions per minute (rpm) to 10000 revolutions per minute (rpm), and the ratio between the mean radius (Rmoy) of the low pressure compressor and the inner radius (Ric) of the rotor of the first stage having a value in the range 1.1 to 1.

4.

2. The propulsion system of claim 1, wherein, the diameter (D) of the fan rotor (22) having a value in the range 1.270 meters to 4.300 meters.

3. The propulsion system of claim 2, wherein, the fan section (20) being unducted and the diameter of the fan rotor (22) having a value in the range 1.980 meters to 4.300 meters.

4. The propulsion system of claim 2, wherein, the fan section (20) being ducted and the diameter of the fan rotor (22) having a value in the range 1.270 meters to 3.048 meters.

5. The propulsion system of any one of claims 1 to 4, wherein, the low pressure compressor (4) being a three-stage compressor (44, 46, 48) and the compression ratio of the low pressure compressor having a value in the range 2.8 to 4.8, the compression ratio being defined as the ratio between the gas pressure leaving the low pressure compressor and the gas pressure entering the low pressure compressor.

6. The propulsion system of any one of claims 1 to 4, wherein, the low pressure compressor (4) being a two-stage compressor (44, 46) and the compression ratio of the low pressure compressor having a value in the range 1.7 to 2.5, the compression ratio being defined as the ratio between the gas pressure leaving the low pressure compressor and the gas pressure entering the low pressure compressor.

7. The propulsion system according to any one of the preceding claims, the propulsion system having a bypass ratio having a value in the range 12 to 16, the bypass ratio being defined as the ratio between the mass flow rate of secondary air (F2) flowing through the fan section around the main body and the mass flow rate of primary air (Fl) flowing through the main body.

8. The propulsion system of any of the preceding claims, wherein, the inner radius (Ric) of the rotor (44r) of the first stage (44) having a value in the range 250 millimeters to 350 millimeters.

9. The propulsion system of any of the preceding claims, wherein, the reduction ratio of the reduction mechanism (19) having a value in the range 2.5 to 11.

10. The propulsion system of any of the preceding claims, wherein, The value of the peripheral speed at the tip of the blades (14) of the fan rotor (22) is in the range 270 m / s to 380 m / s.

11. The propulsion system of any of the preceding claims, wherein, The value of the hub-tip ratio of the fan rotor (22), defined as the ratio between the inner radius (Ri) of the fan rotor and the outer radius (Re) of the fan rotor, is in the range 0.22 to 0.

32.

12. The propulsion system of any of the preceding claims, wherein, The value of the ratio between the first chord (C1) measured at the tip of the blades (14) of the fan rotor (22) and the second chord (C2) measured at the blade root (14) of the fan rotor (22) is in the range 0.56 to 0.

80.

13. The propulsion system of any of the preceding claims, wherein, The value of the solidity of the fan rotor (22) at the tip of the blades (14) of the fan rotor (22), defined as the ratio between the value of the first chord (C1) measured at the tip of the blades (14) and the value of the inter-blade spacing (P), is in the range 1.0 to 1.

3.

14. The propulsion system of any of the preceding claims, wherein, The fan rotor (22) has a number of blades in the range 16 to 24.

15. The propulsion system of any of the preceding claims, wherein, The fan section (20) comprises a fan stator (16) having a number of outlet guide vanes (17) in the range 36 to 52.

16. The propulsion system of any of the preceding claims, wherein, The low pressure turbine (8) comprises a number of stages in the range 3 to 5.

17. The propulsion system of any of the preceding claims, wherein, The high pressure turbine (7) comprises two stages.

18. The propulsion system of any of the preceding claims, wherein, The high pressure compressor (5) comprises a number of stages in the range 8 to 11 stages.

19. The propulsion system of any of the preceding claims, wherein, The value of the power density of the low pressure compressor (4), defined by the following equation: where: BPR is the bypass ratio of the propulsion system, and BPR is measured when the propulsion system is stationary at sea level and in standard atmosphere at take-off rating; Net is the number of stages of the compressor; Um is the average peripheral speed of the compressor rotor measured when the propulsion system is stationary at sea level and in standard atmosphere at take-off rating, and is in meters per second (m / s); Rext moy is the average outer radius of the compressor, equal to the arithmetic average of the outer radii of the low pressure compressor rotor, corresponding to the distance between the tip of the rotor and the rotation axis at the intermediate distance between the leading edge and the trailing edge of the rotor at the tip, and is in meters (m).

20. An aircraft (100) comprising a propulsion system (1) according to any one of the preceding claims.

21. A method of designing a propulsion system (1), the propulsion system comprising: - a fan shaft; - a fan section comprising a fan rotor driven in rotation by the fan shaft; - a main body comprising a compressor section, a combustion chamber and a turbine section, the compressor section comprising a low-pressure compressor and a high-pressure compressor, the turbine section comprising a high-pressure turbine and a low-pressure turbine, the high-pressure turbine driving in rotation the high-pressure compressor through a high-pressure shaft, the low-pressure turbine driving in rotation the low-pressure compressor through a low-pressure shaft; - a reduction mechanism coupling the low-pressure shaft and the fan shaft, so as to drive the fan shaft at a lower speed than the rotation speed of the low-pressure shaft; the low-pressure compressor comprising a first stage and at least one other stage, each stage comprising in succession in the direction of the gas flow a rotor and a stator, the method comprising the steps of: - selecting, for the low-pressure compressor, a value of maximum speed in the range 8500 revolutions per minute (rpm) to 10000 revolutions per minute (rpm), and - selecting an average radius (Rmoy) of the low-pressure compressor and an inner radius (Ric) of the rotor, so that the ratio between the average radius (Rmoy) of the low-pressure compressor and the inner radius (Ric) of the rotor of the first stage has a value in the range 1.1 to 1.

4.

22. A method of designing a propulsion system according to any one of claims 2 to 18.

23. A method of designing a propulsion system according to claim 21 or 22, further comprising the step in which, for the low-pressure compressor (4), a value of power density is selected in the range 0.64 to 0.97, the power density of the compressor being defined by the following equation: wherein: BPR is the bypass ratio of the propulsion system, and BPR is measured when the propulsion system is stationary at sea level and in standard atmosphere at take-off rating; Net is the number of stages of the compressor; Um is the average peripheral speed of the compressor rotors measured when the propulsion system is stationary at sea level and in standard atmosphere at take-off rating, and is expressed in meters per second (m / s); Rext_moy is the average outer radius of the compressor, equal to the arithmetic average of the outer radii of the low-pressure compressor rotors, the outer radius corresponding to the distance between the tip of the rotor and the rotation axis at an intermediate distance between the leading edge and the trailing edge of the rotor at the tip, and expressed in meters (m).

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