PROPULSION PROPELLER FOR AN AIRCRAFT TURBOMACHINE

Abstract

Propulsion propeller for an aircraft turbomachine, comprising a plurality of blades (16) each connected to a root (18) which defines a pitch axis (B) of a rotor blade (14), each blade comprising a virtual stacking line (E) passing through centers of gravity of a set of aerodynamic sections or midpoints of a set of chord lines, said stacking line being defined in an orthonormal XYZ reference frame as follows: - for 0% ≤ H ≤ 50%: Y1 ≤ Ys ≤ Y2, - for 50% ≤ H ≤ 75%: Y1 ≤ Ys ≤ +8.4xH - 4.25xY2, - for 75% ≤ H ≤ 100%: 10.8xH - 12.6xY1 ≤ Ys ≤ +8.4xH - 4.25xY2, where H is a stacking line height measured along the Z axis, and Ys is a stacking line offset measured along the Y axis. Figure for abstract: Figure 2

Description

Title of the invention: PROPULSION PROPELLER FOR AN AIRCRAFT TURBOMACHINE Technical field of the invention

[0001] The present invention relates to a propulsion propeller for an aircraft turbomachine, as well as a turbomachine comprising such a propeller. Technological background

[0002] Turbomachines generally comprise a ducted propeller called a fan, or an unducted propeller, equipped with variable-pitch movable rotor blades. A propeller equipped with variable-pitch or variable-pitch blades makes it possible to adjust the pitch or orientation of the blades according to the flight parameters so as to optimize the operation of the propeller.

[0003] The profile of the propeller and more precisely the profile of the rotor blades of the propeller impact the aerodynamic performance of the turbomachine equipped with it. It is particularly well known that any rotating blade subjected to a given aerodynamic force and centrifugal force will deform until it reaches an equilibrium position. It is also known that the average stresses of the blade have a direct influence on its fatigue resistance.

[0004] This is why engine manufacturers are continually working on the geometry of profiles to improve aerodynamic performance while respecting the mechanical constraints of manufacturing and operation.

[0005] Also, an objective of the invention is to optimize the average stress seen by the blading without making any modification to the materials used. Summary of the invention

[0006] There is therefore proposed a propulsion propeller for an aircraft turbomachine, comprising a rotor hub rotatable about a longitudinal axis, and rotor blades mounted on the hub pivotably about pitch axes, each rotor blade comprising a blade connected to a root which defines the pitch axis of the blade, this blade comprising a lower surface and an upper surface which extend from a leading edge to a trailing edge, this blade further having a plurality of aerodynamic sections and a plurality of chord lines, each chord line connecting the leading edge to the trailing edge in a rectilinear manner for a given aerodynamic section, the plurality of aerodynamic sections and the plurality of chord lines being perpendicular to the pitch axis and being respectively stacked from a first radial end of the blade located on the root side to a second opposite radial end of the blade,the blade further comprising a virtual line, stacking line passing through centers of gravity of all aerodynamic sections or passing through midpoints of all chord lines, in which, in an orthonormal XYZ reference frame which has as its center a point of intersection between the pitch axis and the first end of the blade, and in which the X axis of the reference frame is parallel to the longitudinal axis, the Z axis of the reference frame coincides with the pitch axis, and the Y axis of the reference frame is perpendicular to the X and Z axes and is oriented from the intrados to the extrados of the blade, the virtual stacking line being defined as follows: for 0% < H < 50%: Y! < Ys < Y2, for 50% < H < 75%: Yj < Ys < +8.4xH - 4.25xY2, for 75% < H < 100%: 10.8xH - 12.6xYj < Ys < +8.4xH - 4.25xY2, in which H is a height of the virtual stacking line measured along the Z axis, and Ys is an offset of the virtual stacking line measured along the Y axis.

[0007] Thus, thanks to the invention, an improvement in the control of the average stresses seen by the blading is ensured. Indeed, when the virtual stacking line is included in the limits defined previously, this allows for a balancing of the stresses between the intrados and the extrados of the rotor blade blade so as to increase the centrifugal return effect which depends in particular on the mass distribution around the pitch axis. This also makes it possible to optimize the service life of the blade.

[0008] The propulsion propeller according to the invention may comprise one or more of the following characteristics, taken in isolation from one another or in combination with one another: - the value Yi is equal to -70 mm and the value Y2 is equal to 80 mm; - the offset of the virtual stacking line evolves, depending on the height of the virtual stacking line, in a strictly monotonous manner; - a projection of the virtual stacking line onto a plane perpendicular to the longitudinal axis has at least one inflection point; - a projection of the virtual stacking line onto a plane perpendicular to the longitudinal axis has at least one local maximum and / or minimum; - the rotor hub has a hub ratio of between 0.25 and 0.30, the hub ratio corresponding to a ratio between a first radius of the hub at a radial axis of a rotor blade and a second radius of the propeller at the second end of a blade of a rotor blade; - at least one blade has an elongation of between 2 and 3.5, preferably between 2.5 and 3, the elongation corresponding to a ratio between a height of the blade and an average chord of the blade along the height of the blade; - at least one rotor blade defines a pitch angle, measured at 75% of a helix radius, which is between 55 degrees and 75 degrees; - the first radial end of at least one blade has a thickness greater than a thickness of the second radial end of the blade in question, so that the thickness of the blade decreases monotonically from the first end towards the second end.

[0009] The invention also relates to an aircraft turbomachine comprising at least one propulsion propeller as described above, this propeller being unducted. Brief description of the figures

[0010] The invention will be better understood with the aid of the following description, given solely by way of example and with reference to the appended drawings in which:

[0011] [Fig. 1] represents a schematic view of the propulsion propeller according to the invention, in particular in the environment of a turbomachine,

[0012] [Fig. 2] represents a schematic view of a rotor blade of the propeller according to the invention in an XZ reference frame,

[0013] [Fig.3] represents a schematic view of the rotor blade of [Fig.2] in a YZ reference frame,

[0014] [Fig.4] represents a schematic view of the rotor blade of [Fig.2] according to section plane II,

[0015] [Fig.5] is a graph showing the shape in a YZ plane of the virtual stacking line as a function of the height,

[0016] [Fig.6] is a graph showing the appearance in a YZ plane of different virtual stacking lines as a function of height,

[0017] [Fig.7] is a graph showing the shape in a YZ plane of different virtual stacking lines as a function of height, according to other embodiments of the invention,

[0018] [Fig. 8] represents a comparison of schematic views of a first aerodynamic section taken at a first radial end of a blade and a second aerodynamic section taken at a second radial end of this same blade. Detailed description of the invention

[0019] [Fig.l] schematically represents a propulsion propeller 10 for an aircraft turbomachine, in particular for example in an environment of a turbomachine 100 which comprises a motor 102 for driving the propeller 10. This motor 102 extends for example in an external casing 104 and is located downstream, in the direction of flow of the gases, of the propeller 10. The motor 102 is for example a gas generator conventionally comprising, from upstream to downstream, at least one compressor, at least one compressor. at least one combustion chamber and at least one turbine intended to drive the propeller 10 and the at least one compressor. The turbomachine 100 may further comprise an air inlet 106 for supplying primary flow to the engine 102. This air inlet 106 may be provided downstream of the propeller.

[0020] The propulsion propeller 10 according to the invention comprises a rotor hub 12 which is mobile in rotation about a longitudinal axis A. The propeller 10 also comprises rotor blades 14 mounted on the hub 12 in a pivoting manner about setting axes B.

[0021] With reference to Figures 2 to 4, each rotor blade 14 further comprises a blade 16 connected to a root 18 which defines the pitch axis B of the blade 14. This blade 16 further comprises a lower surface 20 and an upper surface 22 which extend from a leading edge 24 to a trailing edge 26. The lower surface 20 may be concave. The upper surface 24 may be convex. The leading edge 24 makes it possible to separate the lower surface 20 from the upper surface 22 in the upstream part of the blade 16, while the trailing edge 26 makes it possible to separate the lower surface 20 from the upper surface 22 in the downstream part of the blade 16.

[0022] In the following, an orthonormal XYZ reference frame is defined. This XYZ reference frame has a point O as its origin. The X axis of the reference frame is parallel to the longitudinal axis A, the Z axis of the reference frame is parallel to the setting axis B and the Y axis is perpendicular to the X and Z axes. The Y axis is furthermore oriented from the intrados 20 to the extrados 22 of the blade 16.

[0023] The blade 16 further has a plurality of aerodynamic sections S. An example of an aerodynamic section S is shown in [Fig.4]. Aeronautical section S is understood to mean an area defined by the intrados 20, the extrados 22, the leading edge 24 and the trailing edge 26.

[0024] The blade 16 also has a plurality of chord lines C. By chord line C is meant a virtual line connecting in a rectilinear manner the leading edge 24 to the trailing edge 26 of the blade 16 for a given aerodynamic section S. Such a chord line C is shown in [Fig.4]. It is also understood that there are as many chord lines C as there are aerodynamic sections S.

[0025] The plurality of aerodynamic sections S and the plurality of chord lines C are furthermore perpendicular to the pitch axis B. In other words, a given aerodynamic section S and a given chord line C extend in the same plane perpendicular to the pitch axis B and parallel to the longitudinal axis A of the propeller 10. The plurality of aerodynamic sections S and the plurality of chord lines C are respectively stacked from a first radial end 28 of the blade 16 located on the side of the root 18 to a second opposite radial end 30 of the blade 16.

[0026] The blade 16 comprises a virtual stacking line E passing through centers of gravity CG of all aerodynamic sections S or, alternatively, passing through midpoints PM of all the chord lines C. By midpoint PM, for a given chord line C, is meant the point located midway between the edge leading edge 24 and trailing edge 26; in other words, at 50% of the chord line C.

[0027] In such a case, the orthonormal reference frame XYZ is defined such that the point O is a point of intersection between the setting axis B and the first end 28 of the blade 16, and the axis Z coincides with the setting axis B. In other words, the origin of the reference frame XYZ is located at the external surface of the hub 12.

[0028] [Fig.5] illustrates very schematically and not to scale the limits between which the virtual stacking line E is defined in a YZ plane of the orthonormal reference frame. Thus, the virtual stacking line E is defined as follows:

[0029] - for 0% < H < 50%: Yj < Ys < Y2,

[0030] - for 50% < H < 75%: Yj < Ys < Y4,

[0031] - for 75% < H < 100%: Y3 < Ys < Y4.

[0032] H represents a height of the virtual stacking line E measured along the Z axis. Ys represents an offset of the virtual stacking line E measured along the Y axis.

[0033] This means that for a height H of the stacking line E of between 0 and 50% of this height H, the stacking line E is located at a distance from the Z axis of the XYZ reference frame of between Yi and Y2, this distance being an offset Ys on the Y axis. Advantageously, Yi is equal to -70 mm, preferably equal to -50 mm and even more preferably equal to -30 mm. Advantageously, Y2 is equal to 80 mm, preferably equal to 60 mm and even more preferably equal to 40 mm.

[0034] In other words, for 0% < H < 50%: -70 mm < Ys < +80mm, preferably -50 mm < Ys < +60mm, and even more preferably -30 mm < Ys < +40mm.

[0035] For a height H of the stacking line E comprised between 50 and 75% of this height H, the stacking line E is located at a distance from the Z axis of the XYZ reference frame comprised between Yi and Y4, this distance being an offset Ys on the Y axis. Advantageously, Y4 is equal to 8.4xH - 4.25xY2, preferably equal to 8.4xH - 4.5xY2 and even more preferably equal to 8.4xH - 4.75Y2.

[0036] In other words, for 50% < H < 75%: -70 mm < Ys < +8.4xH - 4.25xY2, preferably -50 mm < Ys < +8.4xH - 4.5xY2, and even more preferably -30 mm < Ys < +8.4xH - 4.75xY2.

[0037] For a height H of the stacking line E between 75 and 100% of this height H, the stacking line E is located at a distance from the Z axis of the XYZ reference frame between Y3 and Y4, this distance being an offset Ys on the Y axis. Advantageously, Y3 is equal to 10.8xH - 12.6xYi even more preferably equal to 10.8xH - 12.3xYb

[0038] In other words, for 75% < H < 100%: 10.8xH - 12.6xYi < Ys < +8.4xH - 4.25xY2, preferably 10.8xH - 12.6xYi < Ys < +8.4xH - 4.5xY2, and even more preferably 10.8xH - 12.3xYi < Ys < +8.4xH - 4.75xY2.

[0039] [Fig.6] is similar to [Fig.5] and shows a variant where the virtual stacking line E' is between more restricted limits than those shown in [Fig.5].

[0040] In the particular case where the stacking line E passes through the centers of gravity CG of all the aerodynamic sections S, the displacement of one or more centers of gravity CG of one or more given aerodynamic sections S makes it possible to influence the average stresses of the rotor blade 14. Indeed, this makes it possible to exploit the stresses created by the rotation of the propeller such as the centrifugal forces which are exerted within the blade 16 for example.

[0041] With the limits given previously for the stacking line E, the constraints between the intrados 20 and the extrados 22 of the blade 16 of the rotor blade 14 are balanced so as to increase the centrifugal return effect which depends in particular on the mass distribution around the setting axis B. This also makes it possible to optimize the service life of the blade 16.

[0042] The same effects are obtained when the stacking line E considered passes through the midpoints PM of all the chord lines C.

[0043] [Fig.7] illustrates a projection of the virtual stacking line E, El, E2 onto a plane perpendicular to the longitudinal axis A.

[0044] Preferably, the offset Ys of the virtual stacking line E evolves, as a function of the height H, in a strictly monotonous manner.

[0045] More preferably, the projection of the virtual stacking line E1, E2 onto a plane perpendicular to the longitudinal axis A has at least one inflection point PL

[0046] More preferably, the projection of the virtual stacking line E1, E2 onto a plane perpendicular to the longitudinal axis A has at least one local maximum and / or minimum.

[0047] The hub 12 has a hub ratio corresponding to a ratio between a first radius Rei of the hub 12 at a radial axis of a rotor blade 14 and a second radius Re2 of the propeller 10 at the second end 30 of a blade 16 of a rotor blade 14. Advantageously, this hub ratio is between 0.25 and 0.30.

[0048] The blade 16 has an elongation corresponding to a ratio between a total height of the blade 16 and an average chord of this blade 16 along this height. The total height of the blade 16 is defined as being the distance separating the first end 28 from the second end 30 of the blade 16. The average chord is defined as being the integral of the local chord over the total height of the blade:

[0049] [Math.l] = £ J

[0050] Cm being the average chord of the blade, L being the total height of the blade, h being a local height, c being a local chord.

[0051] By local chord is meant the chord obtained at a given local height.

[0052] Advantageously, the blade 16 has an elongation of between 2 and 3.5. Preferably, the blade 16 has an elongation of between 2.5 and 3. This allows for better balance of the masses on the blade along the X axis.

[0053] With reference to [Fig.8], conventionally, for each aerodynamic section SI, S2, a skeleton curve SK can be defined as being the curve halfway between the intrados 20 and the extrados 22 and which extends from the leading edge 24 towards the trailing edge 26. This skeleton curve SK is then the line connecting the positions of the centers of a plurality of inscribed circles Cn, in the aerodynamic section SI, S2.

[0054] The thickness of an aerodynamic section SI, S2 can be defined as being the distance separating the intrados 20 from the extrados 22 in a direction normal to the skeleton curve SK. In other words, with reference to [Fig. 8], the thickness at a point of the aerodynamic section SI, S2 can be substantially equal to the diameter of an inscribed circle Cn, cn. Thus, in the comparative view of [Fig. 8], the first aerodynamic section SI, taken at the first end 28 of the blade 16, has a plurality of inscribed circles C1, C2, Cn each having a diameter Dl, D2, Dn. The second aerodynamic section S2, taken at the second end 30 of the blade 16, has a plurality of inscribed circles c1, c2, cn each with a diameter dl, d2, dn.Each of the diameters Dl, D2, Dn of the inscribed circles Cl, C2, Cn of the first aerodynamic section SI is greater than the diameter dl, d2, dn of the corresponding inscribed circles cl, c2, cn of the second aerodynamic section S2. In other words, the thickness of the first aerodynamic section SI is greater than the thickness of the second aerodynamic section S2. It is then understood that for at least one blade 16, the first radial end 28 has a thickness greater than a thickness of the second radial end 30 of the blade 16 in question, so that, advantageously, the thickness of the blade 16 decreases monotonically from the first radial end 28 towards the second radial end 30.

[0055] Referring again to [Fig. 4], the aerodynamic section S, taken at a given height of the blade 16, comprises a chord line C connecting the leading edge 24 to the trailing edge 26. The orientation of the chord line C can change depending on the height considered. Furthermore, the leading 24 and trailing 26 edges are separated by a distance which can also change depending on the height considered. This has already been understood by considering the virtual stacking line E passing through the midpoints PM of all the chord lines.

[0056] There is thus an angle 0 between any plane P perpendicular to the longitudinal axis A and the chord line C at a given height. This angle 0 can therefore change depending on the height considered. For convenience, this angle 0 is considered to be representative of the pitch angle of the propeller 10. The pitch angle 0 is measured at a radius Re2 of the propeller 10 and is measured in particular at 75% of this radius Re2 relative to the longitudinal axis A, that is to say in the direction of the second radial end 30 of the blade 16.

[0057] Advantageously, at least one rotor blade 14 has a pitch angle 0, measured at 75% of the radius Re2 of the propeller 10, which is between 55 degrees and 75 degrees. This pitch angle 0 is representative of a blade taken in cruising mode.

[0058] The radius Re2 taken at 75% is considered to be representative of a profile loaded from an aerodynamic point of view or capable of contributing significantly to the traction of the propeller 10.

[0059] The propulsion propeller according to the invention which has been described in the above has the advantage of improving the control of the average stresses seen by the blading. Indeed, when the virtual stacking line is included in the limits defined previously, this makes it possible to have a balancing of the stresses between the intrados and the extrados of the blade of the rotor blade so as to increase the centrifugal return effect which depends in particular on the distribution of mass around the pitch axis.

[0060] Another advantage is to optimize the life of the blade of the vane without changing the nature or properties of the materials used for the construction of the vane.

[0061] The invention also relates to an aircraft turbomachine 100 which comprises at least one propulsion propeller 10 as described above. In the turbomachine 100, the propeller 10 may be unducted as shown in [Fig.l].

[0062] Turbomachines equipped with at least one unducted propeller are known by the English term "open rotor" or "unducted fan". In this category of turbomachine, there are those which have two unducted and counter-rotating propellers (known by the English acronym UDF for "Unducted Fan") or those having a single unducted propeller and a rectifier comprising several stator blades (known by the English acronym USF for "Unducted Single Fan"). The propeller or propellers forming the propulsion part can be placed at the rear of the gas generator (or engine) so as to be of the pusher type or at the front of the gas generator so as to be of the tractor type. This makes it possible to increase the bypass ratio very significantly without being penalized by the mass of the casings or nacelles intended to surround the blades of the propeller or fan.

Claims

Claims

1. Propulsion propeller (10) for an aircraft turbomachine, comprising a rotor hub (12) rotatable about a longitudinal axis (A), and rotor blades (14) mounted on said hub pivotably about pitch axes (B), each rotor blade comprising a blade (16) connected to a root (18) which defines the pitch axis (B) of the blade, said blade comprising a lower surface (20) and an upper surface (22) which extend from a leading edge (24) to a trailing edge (26), said blade further having a plurality of aerodynamic sections (S) and a plurality of chord lines (C), each chord line (C) connecting rectilinearly said leading edge (24) to said trailing edge (26) for a given aerodynamic section (S),said plurality of aerodynamic sections (S) and said plurality of chord lines (C) being perpendicular to the pitch axis (B) and being respectively stacked from a first radial end (28) of the blade located on the side of the root (18) to a second opposite radial end (30) of the blade, said blade (16) further comprising a virtual stacking line (E) passing through centers of gravity (CG) of all said aerodynamic sections (S) or passing through midpoints (PM) of all said chord lines (C), characterized in that in an orthonormal XYZ coordinate system which has as its center a point (O) of intersection between the pitch axis (B) and said first end (28) of the blade, and in which the X axis of said orthonormal XYZ coordinate system is parallel to the longitudinal axis (A), the Z axis of said orthonormal XYZ coordinate system coincides with the pitch axis (B),and the Y axis of said orthonormal reference frame XYZ is perpendicular to the X and Z axes and is oriented from the intrados (20) to the extrados (22) of the blade, said virtual stacking line (E) being defined as follows:, for 0% < H < 50%: Y! < Ys < Y2, for 50% < H < 75%: Yj < Ys < +8.4xH - 4.25xY2, for 75% < H < 100%: 10.8xH - 12.6xYj < Ys < +8.4xH -4.25xY2, where H is a height of the virtual stacking line (E) measured along the Z axis, and Ys is an offset of the virtual stacking line measured along the Y axis.

2. A propeller according to claim 1, wherein the value Yi is equal to -70 mm and the value Y2 is equal to +80 mm.

3. Propeller according to any one of claims 1 or 2, in which the offset (Ys) of said virtual stacking line (E) evolves, as a function of the height (H) of said virtual stacking line (E), in a strictly monotonous manner.

4. Propeller (10) according to any one of claims 1 or 2, wherein a projection of said virtual stacking line (E') onto a plane perpendicular (P) to the longitudinal axis (A) has at least one inflection point (PI).

5. Propeller (10) according to any one of claims 1 or 2, wherein a projection of said virtual stacking line (E) onto a plane perpendicular to the longitudinal axis has at least one local maximum and / or minimum.

6. A propeller according to any one of claims 1 to 5, wherein the rotor hub (12) has a hub ratio of between 0.25 and 0.30, said hub ratio corresponding to a ratio between a first radius (ReO) of the hub at a radial axis of a rotor blade (14) and a second radius (Re2) of the propeller (10) at the second end (30) of a blade (16) of a rotor blade (14).

7. Propeller (10) according to any one of claims 1 to 6, wherein at least one blade (16) has an elongation of between 2 and 3.5, preferably between 2.5 and 3, said elongation corresponding to a ratio between a height of the blade and an average chord of said blade along said height of the blade.

8. Propeller (10) according to any one of claims 1 to 7, in which the first radial end (28) of at least one blade (16) has a thickness greater than a thickness of the second radial end (30) of the blade in question, so that the thickness of the blade (16) decreases monotonically from said first end (28) towards said second end (30).

9. A propeller (10) according to any one of claims 1 to 8, wherein at least one rotor blade (14) defines a pitch angle (0), measured at 75% of a radius (Re2) of the propeller (10), which is between 55 degrees and 75 degrees.

10. Aircraft turbomachine (100) comprising at least one propulsion propeller (10) according to any one of claims 1 to 9, in which said propeller (10) is unducted.

Images