Aerial propulsion system with improved propulsion efficiency
By introducing a reduction gear mechanism into the aerospace propulsion system and optimizing the design parameters of the low-pressure compressor, the size constraint problem of the low-pressure compressor was solved, and high efficiency and low noise performance under high bypass ratio were achieved.
Patent Information
- Application Number
- CN202480035097.1
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Priority Date
- 2023-05-26
- Filing Date
- 2024-05-24
- Publication Date
- 2026-01-02
AI Technical Summary
In existing aerospace propulsion systems, the performance of low-pressure compressors is limited by size constraints, especially the space constraints between the low-pressure shaft and the low-pressure compressor, which affects the optimization of overall efficiency and bypass ratio.
By introducing a reduction gear into the aerospace propulsion system, the low-pressure shaft and the fan shaft are connected, allowing the fan shaft to rotate at a speed lower than that of the low-pressure shaft. This optimizes the design parameters of the low-pressure compressor, such as the average radius ratio and aerodynamic load, ensuring that the low-pressure compressor maintains high efficiency under high bypass ratios.
Without increasing the compressor size, the compression ratio of the low-pressure compressor and the thermal efficiency of the propulsion system were improved, noise was reduced, and overall performance was optimized.
Smart Images
Figure CN121263586A_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present disclosure relates generally to the field of propulsion systems, and more particularly to an aeronautical propulsion system comprising a ducted or unducted fan and having a high, even very high, bypass ratio. BACKGROUND
[0002] An aeronautical propulsion system generally comprises, in the direction of the gas flow, from upstream to downstream, a fan section, a compressor section, possibly comprising a low pressure compressor and a high pressure compressor, a combustion chamber and a turbine section, possibly comprising a high pressure turbine and a low pressure turbine.
[0003] When the propulsion system is running, the high pressure compressor is driven in rotation by the high pressure turbine via a high pressure shaft. The fan and the optional low pressure compressor are driven in rotation by the low pressure turbine via a low pressure shaft.
[0004] One of the objectives of the technical research is to improve the environmental performance of the aircraft. This is why, at each stage of the design and development, the relevant factors are taken into account in order to obtain components and aeronautical products that are less energy-intensive, more environmentally friendly, the integration and use of which in civil aviation have moderate environmental consequences, with the ultimate aim of improving the energy efficiency of the aircraft.
[0005] Thus, in order to improve the propulsive efficiency of the propulsion system and to reduce the specific consumption of the propulsion system and the noise emitted by the fan section, a propulsion system has been proposed with a high bypass ratio BPR, which corresponds to the ratio between the flow rate of the secondary gas flow and the flow rate of the primary gas flow. To achieve said bypass ratio, the fan section can be decoupled from the low pressure turbine, so as to optimize their respective rotation speeds independently. The decoupling is generally achieved by means of a reduction mechanism 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 rotation speed than the speed of the low pressure shaft, via the reduction mechanism.
[0006] However, in said system, the overall performance of the low pressure compressor can be limited by dimensional constraints, which can in particular result from the space required to install a bearing in a chamber between the low pressure shaft and the mobile wheels of the low pressure compressor. The volume occupied in this chamber can in particular depend on the choice of the rotation speed of said compressor and on the power chosen to be obtained via the compressor, the combustion chamber and the turbine. The volume of the reduction mechanism also affects the dimensional constraints acting on the efficiency of the low pressure compressor located inside the flow channel and on the upstream side of the chamber. Furthermore, taking into account the volume of the flow channel in the low pressure compressor, which comprises between the inner wall of the flow channel and the outer wall of the flow channel, the dimensional constraints will depend on the available space for integrating the compressor in its environment. Thus, the performance of the low pressure compressor depends on the aerodynamics and the space occupied around the compressor. SUMMARY
[0007] One goal of the application is to optimize the performance of a propulsion system, in particular to increase the efficiency of the low-pressure compressor while maintaining a high bypass ratio.
[0008] To this end, in a first aspect, an aircraft propulsion system is proposed, the aircraft propulsion system comprising: - a fan shaft; - a fan section comprising a fan rotor rotated 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 rotating the high-pressure compressor through a high-pressure shaft, the low-pressure turbine rotating 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 comprises 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 maximum speed of the low-pressure compressor has a value in the range 10700 revolutions per minute (rpm) to 13000 revolutions per minute (rpm), and the ratio between the mean radius Rmoy of the low-pressure compressor and the inner radius of the rotor of the first stage is in the range 1.1 to 1.4.
[0009] The low-pressure compressor thus has an optimized performance, in particular an optimized efficiency. The above-mentioned range regarding the ratio between the mean radius Rmoy of the low-pressure compressor and the inner radius of the rotor of the first stage encompasses high-speed low-pressure compressors. The use of high-speed low-pressure compressors makes it possible to obtain a higher low-pressure compression ratio than in the prior art, without increasing the volume of the compressor. The overall compression ratio of the compressor section and the thermal efficiency of the propulsion system are thus both improved.
[0010] The mean radius (Rmoy) of the compressor is defined as the mean value of the radius of each rotor, corresponding to the distance between the rotation axis X of the propulsion system and a point taken at the middle height and at the middle distance between the leading edge and the trailing edge of the rotor. The inner radius Ric corresponds to the radius of the first rotor, i.e. the most upstream rotor, which is the distance between the rotation axis of the propulsion system and a point taken at the blade root at the leading edge of the rotor of the first stage.
[0011] In particular, in order to have an 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, the load of the compressor being defined by the following equation:
[0012] where: H is the difference in air enthalpy between the outlet and the inlet of the BP compressor and is in Joule per kilogram (J / Kg), N1 is the number of stages of the BP compressor, Umoy is the average tangential velocity at 50% height of the rotor of the BP compressor and is in meters per second (m / s), Umoy can be calculated according to the following equation:
[0013] where Rmoy corresponds to the average radius of the rotor at 50% height, Rcomp is the speed of the compressor in turns per minute (tr / min).
[0014] The aerodynamic load (or load) of the compressor corresponds to the aerodynamic load of the rotor stages of the compressor and corresponds to the ratio between the work demand (compressed air) and the operating characteristics / conditions of the compressor. The higher the value of the load, the greater the compression capacity but with a lower efficiency and greater difficulty of operation of the compressor, in equivalent geometry and same operating conditions. This is why the value of the load of the compressor is preferably not more than 0.33.
[0015] The difference in air enthalpy of the outlet and the inlet of the low pressure compressor corresponds to the difference in enthalpy between the leading edge of the first rotor and the trailing edge of the diffuser (or stator) arranged downstream of the last rotor.
[0016] The tangential velocity of the airflow (Umoy) is defined as the average of the tangential velocity at 50% height of each stage of the low pressure compressor.
[0017] Furthermore, the value of 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.
[0018] Some preferred but non-limiting features of the propulsion system according to the first aspect are as follows, taken individually or in combination: - the value of the diameter (D) of the fan rotor is in the range 1.270 meters (m) to 4.300 m; the diameter is measured at the leading edge of each blade of the fan rotor; with this diameter, the propulsion system can be integrated in a conventional aircraft without having to modify the plane, in particular without having to modify the dimensions of the landing gear or the wings or to reinforce the wings; - the fan section is unducted and the value of the diameter (D) of the fan rotor is in the range 1.980 m to 4.300 m; the 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 1.270 m to 3.048 m; the diameter is measured at the leading edge of each blade of the propeller; the diameter can for example be 2.286 m; - the low pressure compressor is a three stage compressor and the compression ratio of the low pressure compressor is 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; the compression ratio of the BP compressor enables the compression ratio of each stage to be in the range 1.4 to 1.7, enabling the compressor to operate with optimised efficiency; the compression ratio of each stage is defined as the compression ratio of the compressor at one power (1 / n); the compression ratio is measured in particular using a sensor comb in the test engine, in particular by measuring the flow rate and the pressure at 35000 feet of climb, 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 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; the compression ratio of the BP compressor enables the compression ratio of each stage to be in the range 1.4 to 1.7, enabling the compressor to operate with optimised efficiency; - the bypass ratio of the propulsion system is in the range 12 to 16, the bypass ratio being defined as the ratio between the mass flow rate of the secondary flow (F2) flowing through the fan section around the main body and the mass flow rate of the main air (F1) flowing through the main body; this value is preferably applicable to a ducted propeller; - the inner radius of the rotor of the first stage is in the range 250 millimetres (mm) to 350 mm; this range of radius values is in particular valid for engines with short medium range thrust, SMR, in particular from 100000 to 175000 Newton, these values being measured at the propulsion system at standard atmosphere and at sea level, stationary at take-off rating; - the reduction ratio of the reduction mechanism is in the range 2.5 to 11, preferably in the range 2.7 to 6.0, more preferably in the range 2.7 to 3.6; for example, the reduction ratio can be 3.0; - the peripheral speed at the tip of the blades of the fan rotor is in the range 270 m / s to 380 m / s when the propulsion system is stationary at sea level in standard atmosphere at take-off rating; the fan section is preferably ducted; - the hub-tip ratio of the fan rotor is in the range 0.22 to 0.32, preferably in the range 0.235 to 0.300, for example approximately 0.27, the hub-tip ratio being defined as the ratio between the inner radius (Ri) of the fan rotor and the outer radius (Re) of the fan rotor, - the value of 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 0.56 to 0.80; - the value of the solidity of the fan rotor at the tip of the fan rotor blade is in the range 1.0 to 1.3, the solidity being defined as the ratio between the value of the first chord measured at the tip of the blade and the value of the blade-to-blade spacing; - the fan rotor has a number of blades in the range 16 to 24, for example 22; in particular, the number of fan blades is optimized by taking into account the fan diameter, the solidity at the tip of the fan rotor blade and the value of the peripheral speed at the tip of the fan rotor blade (in particular the values described previously); - the fan section comprises a fan stator having a number of outlet guide vanes in the range 36 to 52, preferably in the range 38 to 44, for example 40; in particular, this number of fan stator guide vanes is optimized by taking into account the fan diameter, the solidity at the tip of the fan rotor blade and the value of the peripheral speed at the tip of the fan rotor blade (in particular the values described previously); - the low-pressure turbine comprises a number of stages in the range 3 to 5; this number of stages ensures an optimized aerodynamic load of the low-pressure turbine and makes it possible to prevent excessive turbine mass and / or bulk; - the high-pressure turbine comprises two stages; - the high-pressure compressor comprises a number of stages in the range 8 to 11 stages; - the value of the power density of the low-pressure compressor (4) is in the range 0.64 to 0.97, the power density of the compressor being defined by the following equation:
[0019] 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; N et 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 pounds of thrust and in meters per second (m / s); Rext moy is the average outer radius of the compressor, equal to the arithmetic mean of the outer radii of the low-pressure compressor rotor, the outer radius corresponding to the distance between the tip of the rotor and the axis of rotation at the intermediate distance between the leading edge and the trailing edge of the rotor at the tip, and being 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 10,700 rpm to 13,000 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 1The 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 deceleration mechanism with a fixed planetary support frame is schematically shown; - Figure 5 A first example of a planetary reduction mechanism with a rotating planetary support frame 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 2Unlike 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 4The 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 2In 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 0.9 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 10,700 rpm to 13,000 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 design method includes the following steps: - For low-pressure compressors, select a maximum speed value within the range of 10700 rpm to 13000 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] Engine 1 has a low-pressure compressor with a relatively low compression ratio of 2.01, and therefore a low peripheral speed. Therefore, in order to obtain a good overall compression ratio in the compression range, the compression ratio of the high-pressure compressor must be around 20. A disadvantage of this engine is that if an increase in the overall compression ratio is desired, either an additional stage must be added to the high-pressure compressor, or its average radius must be increased, thereby increasing the overall size of the high-pressure compressor.
[0107] Advantageously, engine 2 has a low-pressure compressor with a compression ratio of 3.01, and therefore a higher peripheral speed than engine 1, thus taking into account both the aerodynamic load of the compressor and maintaining an overall size similar to engine 1. Therefore, it is possible to both reduce the compression ratio of the high-pressure compressor to benefit its aerodynamic design and increase the overall compression ratio of the compression section, thereby improving the engine's thermal efficiency.
Claims
1. An aircraft propulsion system (1), comprising: - Fan shaft (20); - Fan section (2), the fan section includes a fan rotor (22) that rotates from the fan shaft; - Main body (3), the main body includes compressor section (29), combustion chamber (6) and turbine section (30), the compressor section includes low-pressure compressor (4) and high-pressure compressor (5), the turbine section includes high-pressure turbine (7) and low-pressure turbine (8), the high-pressure turbine rotates the high-pressure compressor (5) through high-pressure shaft (10), and the low-pressure turbine rotates the low-pressure compressor (4) through low-pressure shaft (11); - A speed reduction mechanism (19) connects the low-pressure shaft (8) and the fan shaft (20) to drive the fan shaft (20) at a speed lower than that of the low-pressure shaft (8). The low-pressure compressor (4) includes a first stage (44) and at least one other stage (46, 48), each stage including a rotor (44r, 46r, 48r) and a stator (44s, 46s, 48s) in sequence in the direction of gas flow. The maximum speed of the low-pressure compressor (4) is in the range of 10,700 rpm to 13,000 rpm, and The ratio between the average radius (Rmoy) of the low-pressure compressor and the inner radius (Ric) of the rotor of the first stage is in the range of 1.1 to 1.
4.
2. The propulsion system according to claim 1, wherein, The diameter (D) of the fan rotor (22) is in the range of 1.270 meters to 4.300 meters.
3. The propulsion system according to claim 2, wherein, The fan section (20) is ductless, and the diameter of the fan rotor (22) is in the range of 1.980 meters to 4.300 meters.
4. The propulsion system according to claim 2, wherein, The fan section (20) is ducted, and the diameter of the fan rotor (22) is in the range of 1.270 meters to 3.048 meters.
5. The propulsion system according to any one of claims 1 to 4, wherein, The low-pressure compressor (4) is a three-stage compressor (44, 46, 48), and the compression ratio of the low-pressure compressor is in the range of 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 according to any one of claims 1 to 4, wherein, The low-pressure compressor (4) is a two-stage compressor (44, 46), and the compression ratio of the low-pressure compressor is in the range of 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, wherein the bypass ratio of the propulsion system is in the range of 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 (F1) flowing through the main body.
8. The propulsion system according to any one of the preceding claims, wherein, The inner radius (Ric) of the rotor (44r) of the first stage (44) is in the range of 250 mm to 350 mm.
9. The propulsion system according to any one of the preceding claims, wherein, The reduction ratio of the reduction mechanism (19) is in the range of 2.5 to 11.
10. The propulsion system according to any one of the preceding claims, wherein, The value of the outer circumferential velocity at the tip of the blade (14) of the fan rotor (22) is in the range of 270 m / s to 380 m / s.
11. The propulsion system according to any one of the preceding claims, wherein, The hub-to-tip ratio of the fan rotor (22) is in the range of 0.22 to 0.32, and the hub-to-tip ratio is defined as the ratio between the inner radius (Ri) and the outer radius (Re) of the fan rotor.
12. The propulsion system according to any one of the preceding claims, wherein, The ratio between the first chord (C1) measured at the tip of the blade (14) of the fan rotor (22) and the second chord (C2) measured at the root (14) of the blade of the fan rotor (22) is in the range of 0.56 to 0.
80.
13. The propulsion system according to any one of the preceding claims, wherein, The consistency value of the fan rotor (22) at the tip of the blade (14) 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 tip of the blade (14) and the value of the blade spacing (P).
14. The propulsion system according to any one of the preceding claims, wherein, The fan rotor (22) has a plurality of blades, the number of which is in the range of 16 to 24.
15. The propulsion system according to any one of the preceding claims, wherein, The fan section (20) includes a fan stator (16) having a plurality of outlet guide vanes (17) in the range of 36 to 52.
16. The propulsion system according to any one of the preceding claims, wherein, The low-pressure turbine (8) comprises multiple stages, the number of which is in the range of 3 to 5.
17. The propulsion system according to any one of the preceding claims, wherein, The high-pressure turbine (7) comprises two stages.
18. The propulsion system according to any one of the preceding claims, wherein, The high-pressure compressor (5) includes multiple stages, the number of which ranges from 8 to 11.
19. The propulsion system according to any one of the preceding claims, wherein, 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: 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; Net is the number of stages of 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 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 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).
20. An aircraft (100) comprising a propulsion system (1) according to any one of the preceding claims.
21. A design method for a propulsion system (1), the propulsion system comprising: - Fan shaft; - A fan section, the fan section including a fan rotor driven to rotate by the 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, and the low-pressure turbine drives the low-pressure compressor to rotate through a low-pressure shaft; - A reduction gear mechanism that connects the low-pressure shaft and the fan shaft to drive the fan shaft at a speed lower than the rotational speed of the low-pressure shaft; The low-pressure compressor includes a first stage and at least one other stage, each stage including a rotor and a stator in sequence in the direction of gas flow. The method includes the following steps: - For the aforementioned low-pressure compressor, the maximum speed value is selected within the range of 10,700 rpm to 13,000 rpm, and - Select the average radius (Rmoy) of the low-pressure compressor and the inner radius (Ric) of the rotor such 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 is in the range of 1.1 to 1.
4.
22. A method for designing a propulsion system, said propulsion system being the propulsion system according to any one of claims 2 to 18.
23. The design method for the propulsion system according to claim 21 or 22 further comprises the step of, 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, the power density of the compressor being defined by the following equation: in: 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; Net is the number of stages of the compressor; Um is the average circumferential velocity of the compressor rotor measured when the propulsion system is stationary at takeoff rating in standard atmosphere and at sea level, 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 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).