WO2024224017A1 - Variable-pitch vane for an unducted aeronautical thruster - Google Patents

Abstract

The invention relates to a variable pitch vane for an unducted aeronautical thruster, the vane comprising an airfoil (14) which defines a leading edge (18), a trailing edge (20), a pressure side surface (22) and a suction side surface (24), comprising: a mean camber line (LS), a leading edge thickness (Ep0.2) defined as a length of a first segment (S1) intersecting the mean camber line at a first point (A1) located at 0.2% of the total length of the mean camber line and a capture zone thickness (Ep5) defined as a length of a second segment (S2) intersecting the mean camber line at a second point (A2) located at 5% of the total length of the mean camber line. For each sectional plane (P), the ratio (R) of the capture zone thickness (Ep5) to the leading edge thickness (Ep0, 2) is between 2.5 and 8.

Description

DESCRIPTION

TITLE: Variable-pitch blade for an unducted aeronautical propeller

Technical field of the invention

The invention relates to a variable-pitch blade for an unducted aeronautical propeller, as well as a turbomachine comprising such blades.

State of the prior art

The search for minimizing pollutant emissions linked to air transport involves in particular improving the efficiency of propulsion systems, and more particularly the propulsive efficiency which characterizes the efficiency with which the energy which is communicated to the air passing through the engine is converted into thrust force useful for propulsion.

A known principle for improving propulsive efficiency is to modify the elements of the low-pressure system of the thrusters, which contribute immediately to the generation of thrust, in combination with other known elements of the turbomachine, such as the high-pressure body and the combustion chamber. These elements typically include a low-pressure turbine, a low-pressure transmission system driving a fan, a secondary flow straightener guiding the flow of the latter. One solution aims to reduce the compression ratio of the fan, thereby reducing the flow velocity at the engine outlet and the kinetic energy losses associated with it.

One of the main consequences of this reduction in the flow rate at the engine outlet is that it is necessary to pass a greater mass flow of air through the low pressure part (or secondary flow) in order to ensure a given thrust level. This therefore leads to an increase in the engine dilution ratio (or BPR, for ByPass Ratio), defined as the ratio between the mass flow passing through the secondary flow (cold flow), and the mass flow passing through the primary flow (hot flow) and supplying in particular the combustion chamber.

The very large fan diameters required for such dilution would result in a very significant increase in the dimensions of the casing and the nacelle, leading to the development of unducted turbomachines to overcome this problem.

Figure 1 illustrates such an unducted turbomachine 1, of the type designated by the acronym USF (for Unducted Single Fan in English). The turbomachine 1 comprises a propeller wheel 3 upstream (or rotor 3) with variable timing and a downstream rectifier wheel 5 (or stator 5) with fixed or variable timing.

The terms “upstream” and “downstream” are understood relative to a main axis X of the turbomachine 1, which coincides with the axis of rotation of the rotor 3, and relative to a normal flow direction of the air during operation of the turbomachine 1.

The turbomachine 1 is in a “puller” type configuration, that is to say with the rotor 3 and the stator 5 upstream of the turbine, and is generally mounted on the fuselage or under a wing of the aircraft by a mast 7. Alternatively, a turbomachine can be in a so-called “pusher” configuration with the propellers downstream of the turbine for mounting at the rear of the aircraft.

The rotor 3 and the stator 5 each comprise a plurality of blades 10 distributed circumferentially around the main axis X.

The rotor 3 is a movable propeller rotating around the X axis so as to drive the air and generate a primary flow Fp sent into the turbomachine and a secondary flow Fs flowing outside the turbomachine.

This rotor 3 is a variable-pitch propeller with a slow rotation speed, that is to say that each blade 10 comprises a blade mounted so as to rotate about a radial axis, so as to modify its pitch angle to maximize its thrust according to the flight point (takeoff, cruise, landing, etc.), the slow rotation speed making it possible to maximize the propulsion energy efficiency. Such an unducted turbomachine does not comprise an external casing surrounding the secondary flow Fs. Only the primary flow Fp is guided in a central casing of the turbomachine.

The acoustic attenuation of the noise generated by the turbomachine is significantly reduced, particularly for low frequencies (below 500 Hz). The admissible sound levels are highly constrained, particularly for the take-off and landing phases, which implies an optimization of the blades to reduce the noise level, not having access to the classic attenuation solutions of ducted turbomachines.

Finally, the blades of the variable-pitch vanes, in an unducted turbomachine, are particularly sensitive to non-uniform air flows, with non-axial incidence, i.e. forming a non-zero angle with the direction of the main axis X of the turbomachine, due to the absence of an external casing guiding the secondary flow. Such air flows appear notably during the take-off and landing phases of the aircraft and can generate transverse forces and moments in the plane of the propeller. These forces are transient, and vary for each blade during its rotation, with differences depending on the rising or falling position of the blade. Consequently, over one revolution of the engine, the same propeller blade is subjected to variable forces dependent on its azimuthal position. The stator blades located downstream of the propeller will also have a variable incidence depending on their azimuthal position.

A known solution for reducing the noise level generated by the blades is to uniformly reduce the radial length of the blades of the downstream impeller, i.e. of the stator 5 in the case shown. In this way, the impact of the vortices formed at the radially external ends of the blades of the rotor 3 on the stator blades is limited in that these vortices pass radially outside the stator blades. This solution is called "clipping", or "cropping", or "truncation", or even "clipping", of the blades of the downstream impeller. A "clipping" or clipping rate can be defined as the ratio of the difference in radius between the rotor and the stator to the radius of the rotor, generally expressed as a percentage.

However, this solution can still be improved. Indeed, a high clipping rate significantly reduces the propulsive efficiency of the turbomachine. In addition, noise reduction is mainly effective at zero incidence, and does not necessarily provide a satisfactory result in situations with high incidence.

Presentation of the invention

The invention aims to overcome these drawbacks by proposing an unducted turbomachine which is robust to variations in the incidence of air flows at different operating points corresponding to different phases of flight, and which offers satisfactory aerodynamic and acoustic behavior for a wide range of rotation speeds.

For this purpose, the invention relates to a variable-pitch blade for an unducted aeronautical propeller, comprising a blade extending along a blade axis, from a root to a tip of the blade, the blade defining a leading edge, a trailing edge, and intrados and extrados surfaces extending from the leading edge to the trailing edge, the blade comprising, in any section plane orthogonal to the blade axis:

- a skeleton line extending from the leading edge to the trailing edge, equidistant from the intrados surface and the extrados surface, having a total length measured along the skeleton line from the leading edge to the trailing edge,

- a leading edge thickness, defined as a length of a first segment extending from the intrados edge to the extrados edge and intersecting perpendicularly the skeleton line at a first point on the skeleton line located at a distance from the leading edge, measured along the skeleton line, equal to 0.2% of the total length of the skeleton line,

- a capture zone thickness, defined as a length of a second segment extending from the intrados edge to the extrados edge and perpendicularly intersecting the line of skeleton at a second point on the skeleton line located at a distance from the leading edge, measured along the skeleton line, equal to 5% of the total length of the skeleton line, characterized in that, for each section plane orthogonal to the blade axis, the ratio between the thickness of the capture zone and the thickness of the leading edge is between 2.5 and 8.

Such a blade makes it possible to significantly limit aerodynamic separations at the leading edge of the propellers when they operate at low speed and with a high incidence to reach the target thrust during the take-off phase of the aircraft. This then significantly reduces the formation of a downstream vortex, which otherwise constitutes a very energy-consuming propeller wake and is very penalizing from an acoustic point of view.

For each section plane orthogonal to the blade axis, the ratio between the capture zone thickness and the leading edge thickness can be between 3.5 and 5.

Such a characteristic allows a better compromise between the robustness of the blade to variations in incidence and aerodynamic efficiency.

On a lower portion of the blade extending from the root over a height of between 0% and 35% of a total height of the blade measured between the root and the tip, for each section plane of said lower portion, the ratio between the thickness of the capture zone and the thickness of the leading edge may be between 3 and 8.

Such a feature makes it possible to improve robustness to variations in incidence, particularly on the bottom of the blade, which generates the primary flow supplying the turbine in the case of the rotor, and therefore requires less performance from the blade regardless of the flight envelope, or in the case of the stator, which takes up the most thrust.

On an upper portion of the blade extending to the tip, over a height of between 35% and 100% of a total height of the blade measured between the root and the tip, for each section plane of said upper portion, the ratio between the thickness of the capture zone and the thickness of the leading edge may be between 2.5 and 5.

Such a feature makes it possible to optimize the aerodynamic performance of the blade on the upper part while maintaining sufficient robustness to variations in incidence and satisfactory acoustic performance.

The ratio of the capture zone thickness to the leading edge thickness for any section plane of a lower portion of the blade extending from the root over a height of between 0% and 35% of a total height of the blade measured between the root and the tip may be greater than or equal to the ratio of the capture zone thickness to the leading edge thickness for any section plane of an upper portion of the blade extending to the tip over a height of between 35% and 100% of the total height of the blade. Such a feature makes it possible to distribute the aerodynamic performance of the blade over the most critical parts and to improve the robustness to incidence of the lower parts of the blade. Such a variation in the aerodynamic performance of the blade is particularly advantageous for a dual-flow thruster, by optimizing the lower part of the blade for good supply of the radially internal primary flow and by making it possible to have different characteristics for the upper part of the blade, more suited to the supply of the radially external secondary flow.

The blade may comprise, in each section plane orthogonal to the blade axis, a maximum thickness, defined as a length of a third segment extending from the intrados edge to the extrados edge and perpendicularly intersecting the skeleton line at a third point, for which the length of the third segment is maximum over an extent of the skeleton line, in which said third point is located at a distance from the leading edge, measured along the skeleton line, greater than or equal to 15%, and advantageously between 15% and 40% of the total length of the skeleton line.

Such a feature allows to move the maximum thickness away from the leading edge and thus improve the performance of the blade.

The ratio may be strictly increasing from the leading edge to said third point and strictly decreasing from said third point to the trailing edge.

The blade may be a rotor blade mounted on a mobile disk rotating around a main axis.

The invention also relates to an unducted propeller for an aircraft, comprising at least one rotor and one stator, spaced along a main axis of the propeller, at least one of the rotor and the stator comprising a plurality of blades as above, distributed circumferentially around the main axis, in particular between 3 and 25 blades, advantageously between 8 and 16 blades.

Such a number of blades constitutes an advantageous compromise between propulsion energy performance and the noise generated.

Among the rotor and the stator, the one arranged upstream relative to the main axis may comprise at least two more blades than the one arranged downstream.

Such a feature helps to reduce the noise of the turbomachine. Indeed, in the case where the number of rotor and stator blades are equal, the rotor wake assembly interacts with the stator blades simultaneously, which increases the noise levels.

The stator can be arranged downstream and have a clipping rate of between 5% and 15% and in particular between 7% and 12%.

Such a feature allows a further reduction of the noise generated at the stator level, without significantly reducing the propulsive efficiency of the turbomachine. The lengths of the stator blades may be non-uniform, with blade lengths in the lower part of the stator being less than the blade lengths in the upper part of the stator.

Such a feature allows to have a higher clipping rate below the stator, where the noise generated is the most important and therefore where noise reduction is most necessary, and a lower clipping in the upper part, where it is less required. Thus, the compromise between noise reduction and propulsive efficiency is improved.

Each blade may have a chord length defined as the maximum over a span of the blade of a distance between the leading edge and the trailing edge in a cross-section plane to the blade axis, the rotor and the stator having spacings separating neighboring blades, measured in a circumferential direction, in which a solidity of the rotor and the stator, defined as the ratio of the chord length to the spacing separating neighboring blades, may be less than or equal to 3, and in particular less than or equal to 1 for that arranged furthest upstream of the rotor and the stator, relative to the main axis.

A ratio between an axial distance separating the rotor and the stator and an external diameter of the rotor may be between 0.01 and 0.5, preferably between 0.15 and 0.35.

Such a feature allows efficient rectification of the flow at the rotor outlet by the stator and improves the aerodynamic performance of the propeller.

Brief description of the figures

Figure 1 is a side view of an unducted turbomachine according to the invention, Figure 2 is a side view of a rotor blade of the turbomachine of Figure 1, Figure 3 is a cross-sectional view of the blade of Figure 2, and Figure 4 is a cross-sectional profile of the blade of Figures 2 and 3.

Detailed description of the invention

An unducted turbomachine 1 is shown in Figure 1, defining a central axis X and comprising a rotor 3 and a stator 5 spaced along the main axis X. The stator 5 is positioned downstream of the rotor 3.

The turbomachine 1 also comprises at least one engine arranged in its internal space, said engine possibly being a heat engine, in particular of the turboshaft, turbojet, turbofan type, and/or an electric engine, and/or a hydrogen engine, and/or a hybrid engine combining several of these technologies. The rotor 3 and the stator 5 both comprise a plurality of blades 10 extending substantially radially from the main axis X and regularly distributed circumferentially around the axis X. The blades 10 of the rotor 3 may have different dimensions from the blades 10 of the stator 5, in particular different blade lengths.

Advantageously, the rotor 3 comprises at least two blades 10 more than the stator 5.

The rotor 3 has a radius R1, measured from the central axis to a tip of each blade 10 of the rotor 3.

Similarly, the stator 5 has a radius R2 measured from the central axis to a tip of each blade 10 of the stator 5.

According to one embodiment, the radius R2 of the stator 5 is less than the radius R1 of the rotor 3, in particular less than the radius R1 by between 5% and 15% of the value of R1 and in particular by between 7% and 12% of the value of R1. In other words, the stator 5 has a clipping rate, as defined above, of between 5% and 15% and in particular between 7% and 12%. This clipping rate allows an additional reduction in the noise generated at the stator, but remains sufficiently moderate so as not to significantly reduce the propulsive efficiency of the turbomachine.

We define a diameter D of the turbomachine 1 as twice the largest of the two radii R1, R2.

The blades 10 are variable-pitch blades, that is to say, the blades of which are movable in rotation around a radial axis in order to vary the pitch angle of each blade of the rotor 3 or the stator 5 in a controlled manner.

The axes of rotation of the blades, or blade axes, of the rotor blades 3 are included in a plane P1 perpendicular to the main axis X, and the axes of the blades of the stator 5 are included in a plane P2 perpendicular to the main axis X and spaced from the plane P1 by a distance S measured along the main axis X.

In the case where the blades of the stator 5 are fixed-pitch blades, the plane P2 is defined at the level of the center of gravity of the blades 10.

Advantageously, a ratio S/D, between the axial distance S separating the planes P1 and P2, and one separating the rotor 3 from the stator 5, and the external diameter of the turbomachine 1 is between 0.01 and 0.5, preferably between 0.15 and 0.35.

A blade 10 is shown in more detail in Figure 2. The blade 10 comprises a root 12, a blade 14 and a tip 16. The blade 14 extends along a blade axis Z perpendicular to the main axis X and included in the plane P1 described above.

The blade 14 defines a leading edge 18 formed by the most upstream line of the blade 14 and a trailing edge 20 formed by the most downstream line. The blade 14 comprises a lower surface 22 and an upper surface 24 extending from the leading edge to the trailing edge on either side of the blade 14.

A blade height H is defined as the distance measured along the blade axis Z separating the root 12 from the tip 16, and a chord length C as the distance separating the leading edge 18 from the trailing edge 20, measured in a plane of transverse section P.

We define a so-called solidity ratio TT = C/E of the rotor 3 or the stator 5 as the ratio between the chord length C measured at the tip 16 and a spacing E between the tips 16 of two neighboring blades 10 of the rotor 3 or the stator 5.

Advantageously, the solidity is less than 3 for the rotor and the stator, and preferably less than 1 for the one placed furthest upstream, that is to say the rotor 3 in the case shown. A height h of a plane of transverse section P is defined as the distance measured along the blade axis Z between the root 12 and the plane of section P.

Figures 3 and 4 represent transverse sections of the blade 14 in a section plane P located at a height h.

In the plane P, a skeleton line LS is defined as the curved line extending from the leading edge 18 to the trailing edge 20 and equidistant from the intrados surface 22 and the extrados surface 24.

The length of the skeleton line LS is strictly greater than the chord length C in the plane P, which separates the leading edge 18 from the trailing edge 20, measured in a straight line.

The pitch angle y of the blade 14 is represented as the angle between the chord C and the transverse plane P1. The pitch angle y can be modified by rotating the blade 14 around the blade axis Z.

As shown in Figure 4, a thickness of the blade 14 in the section plane P at a height h is defined as the length of a segment perpendicular to the skeleton line LS and extending from the intrados edge 22 to the extrados edge 24.

We thus define a leading edge thickness Ep0.2 as the length of a first segment S1 which intersects the skeleton line LS perpendicularly at a first point A1 located at a distance, measured along the skeleton line LS, equal to 0.2% of the total length of the skeleton line LS.

Similarly, a capture zone thickness Ep5 is defined as the length of a second segment S2 which intersects the skeleton line LS perpendicularly at a second point A2 located at a distance, measured along the skeleton line LS, equal to 5% of the total length of the skeleton line LS.

The ratio R = Ep5/Ep0.2 between the thickness of the capture zone Ep5 and the thickness of the leading edge Ep0.2 is characteristic of the performance of the blade and its robustness to variations in incidence. In particular, low values of R correspond to blades with high aerodynamic performance, while high R values correspond to blades that are robust at incidence.

Thus, R values between 2.5 and 8 provide an advantageous compromise allowing good performance for the blade while reducing the risks of detachment and having satisfactory acoustic performance.

These minimum and maximum values of the ratio R are valid for any section plane P of the blade 14, that is to say over the entire height H of the blade 14.

The behavior of blade 14 is even more satisfactory for R values between 3.5 and 5 over the entire height H.

According to an advantageous embodiment, a lower part of the blade 14 is distinguished, for heights h between 0 and 35% of the total height H, and an upper part of the blade, for heights h between 35% and 100% of the total height H of the blade 14.

Over the entire lower part of the blade 14, the ratio R is advantageously between 3 and 8, emphasizing the resistance of the blade to variations in incidence. Indeed, the aerodynamic performances are less critical in the lower part, which concentrates the primary flow Fp towards the turbine.

Over the entire upper part of the blade 14, the ratio R is advantageously between 2.5 and 5 in order to emphasize the aerodynamic performance of the blade 14 on the upper part generating the secondary flow Fs.

Furthermore, the ratio R over the entire lower part is advantageously greater than the ratio R over the entire upper part of the blade 14.

The application of these criteria is valid regardless of the shape of the leading edge area, whether for a substantially circular leading edge or for a substantially asymmetrical leading edge.

As before, we define a maximum thickness Epmax as the length of a third segment S3 which intersects the skeleton line LS perpendicularly at a third point A3 and for which the measured length is maximum over the entire skeleton line.

Advantageously, the blade thickness varies monotonically from the leading edge to the maximum Epmax, then again monotonically from the maximum Epmax to the trailing edge.

Preferably, the third point A3 is located at a distance from the leading edge measured along the skeleton line LS greater than or equal to 15% of the total length of said skeleton line LS. This makes it possible to move the maximum thickness sufficiently away from the leading edge to have satisfactory performance. More preferably, the point A3 is located at a distance from the leading edge of between 15% and 40% of the total length of said skeleton line LS.

Claims

1. Variable-pitch blade (10) for an unducted aeronautical propeller (1), comprising a blade (14) extending along a blade axis (Z), from a root (12) to a tip (16) of the blade (14), the blade (14) defining a leading edge (18), a trailing edge (20), and intrados (22) and extrados (24) surfaces extending from the leading edge (18) to the trailing edge (20), the blade (14) comprising, in any section plane (P) orthogonal to the blade axis (Z): - a skeleton line (LS) extending from the leading edge (18) to the trailing edge (20), equidistant from the intrados surface (22) and the extrados surface (24), having a total length measured along the skeleton line (LS) from the leading edge (18) to the trailing edge (20), - a leading edge thickness (Ep0.2), defined as a length of a first segment (S1) extending from the intrados edge (22) to the extrados edge (24) and perpendicularly intersecting the skeleton line (LS) at a first point (A1) of the skeleton line (LS) located at a distance from the leading edge (18), measured along the skeleton line (LS), equal to 0.2% of the total length of the skeleton line (LS), - a capture zone thickness (Ep5), defined as a length of a second segment (S2) extending from the intrados edge (22) to the extrados edge (24) and perpendicularly intersecting the skeleton line (LS) at a second point (A2) of the skeleton line (LS) located at a distance from the leading edge (18), measured along the skeleton line (LS), equal to 5% of the total length of the skeleton line (LS), characterized in that, for each section plane (P) orthogonal to the blade axis (Z), the ratio (R) between the capture zone thickness (Ep5) and the leading edge thickness (Ep0.2) is between 2.5 and 8, in which the ratio (R) between the capture zone thickness (Ep5) and the leading edge thickness (Ep0.2) for any section plane (P) of a lower portion of the blade (14) extending from the root (12) over a height (h) between 0% and 35% of a total height (H) of the blade (14) measured between the root (12) and the tip (16) is greater than or equal to the ratio between the thickness of the capture zone (Ep5) and the thickness of the leading edge (Ep0.2) for any section plane (P) of an upper portion of the blade (14) extending to the tip (16) over a height (h) between 35% and 100% of the total height of the blade (14). 2. Blade (10) according to claim 1, in which, for each section plane (P) orthogonal to the blade axis (Z), the ratio (R) between the thickness of the capture zone (Ep5) and the thickness of the leading edge (Ep0.2) is between 3.5 and 5. 3. Blade (10) according to claim 1 or 2, in which, on a lower portion of the blade (14) extending from the root (12) over a height (h) of between 0% and 35% of a total height (H) of the blade (14) measured between the root (12) and the tip (16), for each section plane (P) of said lower portion, the ratio (R) between the thickness of the capture zone (Ep5) and the thickness of the leading edge (Ep0.2) is between 3 and 8. 4. Blade (10) according to one of the preceding claims, in which, on an upper portion of the blade (14) extending to the tip (16), over a height (h) of between 35% and 100% of a total height (H) of the blade (14) measured between the root (12) and the tip (16), for each section plane (P) of said upper portion, the ratio (R) between the thickness of the capture zone (Ep5) and the thickness of the leading edge (Ep0.2) is between 2.5 and 5. 5. Blade (10) according to one of the preceding claims, in which the blade (14) comprises, in each section plane (P) orthogonal to the blade axis (Z), a maximum thickness (Epmax), defined as a length of a third segment (S3) extending from the intrados edge (22) to the extrados edge (24) and perpendicularly intersecting the skeleton line (LS) at a third point (A3), for which the length of the third segment (S3) is maximum over an extent of the skeleton line (LS), in which said third point (A3) is located at a distance from the leading edge (18), measured along the skeleton line (LS), greater than or equal to 15% of the total length of the skeleton line (LS), and advantageously between 15% and 40% of the total length of the skeleton line (LS). 6. Blade (10) according to one of the preceding claims, in which the blade (10) is a rotor blade (3) mounted on a mobile disk rotating about a main axis (X). 7. Unducted propeller (1) for an aircraft, comprising at least one rotor (3) and one stator (5) spaced along a main axis (X) of the propeller (1), at least one of the rotor (3) and the stator (5) comprising a plurality of blades (10) according to one of the preceding claims, distributed circumferentially around the main axis (X), in particular between 3 and 25 blades (10), advantageously between 8 and 16 blades (10). 8. Thruster (1) according to the preceding claim, in which, among the rotor (3) and the stator (3), the one arranged upstream relative to the main axis (X) comprises at least two blades (10) more than the one arranged downstream. 9. Thruster (1) according to claim 7 or 8, in which the stator (5) is arranged downstream and has a clipping rate of between 5% and 15% and in particular between 7% and 12%. 10. A thruster (1) according to one of claims 7 to 9, wherein each blade (14) has a chord length (C) defined as the maximum over an extent of the blade (14) of a distance between the leading edge (18) and the trailing edge (20) in a section plane (P) transverse to the blade axis (Z), the rotor (3) and the stator (5) having spacings (E) separating neighboring blades (14), measured in a circumferential direction, wherein a solidity (TT) of the rotor (3) and the stator (5), defined as the ratio of the chord length (C) to the spacing (E) separating neighboring blades, is less than or equal to 3, and in particular less than or equal to 1 for the one located furthest upstream of the rotor (3) and the stator (5), relative to the main axis (Z). 11. Thruster (1) according to one of claims 7 to 10, in which a ratio (S/D) between an axial distance (S) separating the rotor (3) and the stator (5) and an external diameter (D) of the thruster (1) is between 0.01 and 0.5, preferably between 0.15 and 0.35.