Rotor blade of aircraft and aircraft provided with such blade

Abstract

The invention relates to a blade (20) for a rotor capable of rotating about an axis of rotation (AX), said blade (20) extending along a pitch axis (AXPAS) from a first end (51) to a second end (52), the blade (20) comprising a blade body (25) having a series of sections (S) substantially perpendicular to the pitch axis (AXPAS) from the first end (51) to the second end (52). The blade (20) has sections along its pitch axis with a chord that increases and then decreases, a relative thickness that decreases from an initial section (SINI) to a final section (SF), a torsional angle that continuously increases and then decreases from an initial section (SO) to the initial section (SINI), and an offset that varies with respect to the pitch axis.

Description

[0001] Cross-reference to related applications

[0002] This application claims the benefit of FR2405108, filed on May 17, 2024, the disclosure of which is incorporated herein by reference in its entirety. Technical Field

[0003] The present invention relates to a rotor blade of an aircraft, and an aircraft having the rotor blade.

[0004] The aircraft may include a rotor with at least one blade. Specifically, the blade is rotatable about the rotor's axis of rotation and its pitch axis. For example, each blade is connected to a pitch control system. This pitch control system may include actuators hinged to the lower plate of a set of swashplates, the upper plate of which is connected to each blade via a corresponding pitch linkage. The pitch control system may also include an upper scissor link connecting the upper plate to the rotor shaft and a lower scissor link connecting the lower plate to a fixed housing. Each blade may also be connected to at least one component sometimes referred to as a "drag damper" or "frequency adapter."

[0005] This type of blade, depending on its span, includes a blade root for attachment to a hub, followed by an aerodynamic body. This body provides most of the blade's lift.

[0006] According to another aspect, a rotorcraft has the advantage of being able to operate at high forward speeds and extremely low forward speeds, or even at zero speed during the hovering phase.

[0007] The geometry of the propeller blades affects the operation of the aircraft during high-speed forward flight and low-speed flight.

[0008] In fact, the higher the forward speed, the greater the forces generated by the blades on the pitch control system, or even on the frequency adapter or drag damper. Manufacturers may seek to minimize these forces.

[0009] In addition, manufacturers seek blades that can achieve good performance at low speeds, especially those that can take off with maximum payload.

[0010] However, these two objectives appear to be contradictory. While it is true that reducing the forces generated at high forward speeds can be achieved by acting on the blade geometry in pitch control systems or even in frequency adapters or drag dampers, this typically compromises performance achieved at low speeds, and vice versa.

[0011] Conventional helicopter rotor blades, referred to as "conventional blades" for convenience, are sized to generate acceptable forces in the pitch control system or even in the frequency adapter or drag damper, while enabling favorable performance at low speeds.

[0012] However, some aircraft can operate at higher stable forward cruise speeds, such as greater than 200 knots instead of the 150 knots typically observed for helicopters. Above 200 knots, conventional rotor blades introduce forces into the pitch control system (actuators, pitch levers, swashplates, scissor linkages, etc.) and into the frequency adapter or drag damper, which can affect their lifespan. Background Technology

[0013] Traditionally, blades are defined by the shape of the aerodynamic profiles of segments of the blade and the positioning of these segments relative to each other. Thus, patent FR3045564 describes a blade for a rotor of a rotorcraft, which extends along the blade axis between the blade initiation point and the blade tip, which can be connected to the rotor hub. The blade includes a shaped portion located between the blade initiation point and the blade tip, which is composed of a series of aerodynamic profiles. The blade tip is located at a reference distance from the axis of rotation equal to the rotor radius R. The chord of the profile of the shaped portion (defined as the distance between the leading and trailing edges of the profile) increases between the initiation point of the shaped portion and a first segment located at a first distance between 0.6*R and 0.9*R from the axis of rotation A, decreasing beyond the first segment. The geometric twist of the blade segment profile decreases between the second segment located at a second distance between 0.3*R and 0.4*R from the axis of rotation and the blade tip. The first torsional gradient between the second segment and the third segment located at a third distance between 0.4*R and 0.6*R from the axis of rotation is between -25° divided by radius R and -4° divided by radius R. The second torsional gradient between the third segment and the fourth segment located at a fourth distance between 0.65*R and 0.85*R from the axis of rotation is between -25° divided by radius R and -4° divided by radius R. The third torsional gradient between the fourth segment and the fifth segment located at a fifth distance between 0.85*R and 0.95*R from the axis of rotation is between -16° divided by radius R and -4° divided by radius R. The fourth torsional gradient between the fifth segment and the blade tip is between -16° divided by radius R and 0° divided by radius R.

[0014] The twist of the blades involves changing the angles of the segments relative to each other. The "twist angle" of a segment is the geometric angle formed between the chord of that segment and the chord of a selected reference segment parallel to that blade. For convenience, a positive angle is assumed below to correspond to the upward tilt of the segment's head relative to the reference segment. The change in the twist angle along the blade's span is known as the "law of torsion."

[0015] Patent FR3045565 describes a rotor blade for a rotorcraft, comprising a shaped portion located between the blade's origin and its tip. The blade tip is located at a reference distance from the rotation axis A equal to the rotor radius R. The chord of the shaped portion's profile increases between the origin of the shaped portion and a first segment located at a first distance between 0.6*R and 0.9*R from the rotation axis, decreasing beyond the first segment. The blade has a forward-swept shape between the origin of the shaped portion and a second segment located at a second distance between 0.5*R and 0.8*R from the rotation axis A, with the leading edge forming a first forward-swept angle α1 between 0° and 10° with the blade axis. The blade has a forward-swept shape between the second segment and a third segment S3 located at a third distance between 0.6*R and 0.95*R from the rotation axis, with the leading edge forming a second forward-swept angle α2 between 1° and 15° with the blade axis. In addition, the blade has a sweeping shape pointing towards the rear of the blade between the third section and the blade tip, and the leading edge forms a third sweep angle α3 between -35° and -15° with the blade axis B.

[0016] Documents EP0565413A1, EP0842846A1 and EP0901961A1 are also known. Summary of the Invention

[0017] The object of the present invention is to provide a blade that limits the force generated at high speeds on the pitch control system or even on at least one frequency adapter or drag damper, while enabling lift to be generated at low speeds and especially during hovering, giving the rotorcraft acceptable performance, and for example having performance at least substantially equivalent to that achieved on conventional aircraft.

[0018] This invention relates to a rotor blade for an aircraft, the blade being rotatable about a rotor axis of rotation and about a pitch axis, the blade extending along the pitch axis from a first end to a second end, the blade comprising a blade body having a blade root along the pitch axis and then a main portion formed by a series of segments substantially perpendicular to the pitch axis, the blade root having an initial segment forming the first end, the main portion extending along the pitch axis from the initial segment to a final segment forming the second end, the final segment being located at a distance from the rotation axis equal to a predetermined rotor radius R, each segment extending along a transverse axis from a leading edge to a trailing edge, the leading and trailing edges being separated by a maximum distance constituting a chord, each segment of the blade body having a geometric twist angle relative to a reference segment located at a distance from the rotation axis equal to 70% of the rotor radius R.

[0019] The blades also have the following characteristics:

[0020] The initial section is set at a distance between 20% and 30% of the rotor radius R from the axis of rotation;

[0021] According to the law of blade span variation of the chord of a segment, the chord of a segment increases from the initial segment to the maximum chord reached in the first segment at a first distance between 75% and 80% of the rotor radius R from the axis of rotation. Then it decreases according to the following law: first it decreases slowly to the second segment, and then decreases rapidly beyond the second segment. The second segment is located at a second distance between 80% and 95% of the rotor radius R from the axis of rotation. The dimensionless average aerodynamic chord relative to the rotor radius R is between 0.05 and 0.08. The dimensionless maximum chord relative to the rotor radius R is greater than the dimensionless average aerodynamic chord relative to the rotor radius R and is between 0.06 and 0.1.

[0022] According to the blade span variation law of relative thickness, the relative thickness of the section decreases from the initial section to the initial section from a relative thickness between 0.25 and 0.70 to a relative thickness between 0.12 and 0.15. The main section has a constant relative thickness or decreases and then becomes constant from the initial section to the middle section. The relative thickness of the main section decreases as it moves away from the middle section until it reaches a relative thickness between 0.07 and 0.08 in the final section. The distance of the middle section from the axis of rotation is between 65% and 85% of the rotor radius R.

[0023] According to the torsional law, the geometric twist angle continuously increases from a minimum negative angle to a maximum positive angle from the initial segment to the initial segment. Then, it decreases with a first gradient between -9° divided by the rotor radius R and -13° divided by the rotor radius R to a third segment located at a distance from the rotation axis equal to or greater than 70% and 80% of the rotor radius R. Then, it reaches a fourth segment at a second gradient equal to or greater than the first gradient, located at a distance from the rotation axis equal to or greater than 88% and 90% of the rotor radius R. That is, it produces a second gradient that is the same as or lower than the first gradient. Finally, it reaches the final segment with a third gradient less than the first gradient. That is, it produces a third gradient that is higher than the first gradient.

[0024] According to the offset law, starting from the leading edge, the offset distance that separates the pitch axis from the quarter chord of each segment decreases from the initial segment to the initial section. This offset distance is constant in the main part of the initial segment to the rupture segment, and then decreases. The distance of the rupture segment from the rotation axis is between 80% and 95% of the rotor radius R. The offset distance can be reduced from the rupture segment according to at least a second-order reduction law, that is, according to a function that also decreases according to the derivative and the second derivative.

[0025] The blades are typically sized relative to the desired rotor radius. This rotor radius is a characteristic usually associated with the blades. The length of the blade root can vary from one rotor to another, thus always having the same main section and the same rotor radius, regardless of the dimensions of the rotor components that support the blades.

[0026] Each section at the blade root may have a thick profile, i.e., a relative thickness greater than 15% of its chord, and each section of the main part may have a thin aerodynamic profile, i.e., a relative thickness less than or equal to 15% of its chord.

[0027] The mean aerodynamic chord is defined based on a weighted average of the squares of the radii:

[0028]

[0029] Where c(r) is the law of blade span variation of the chord of the segment, r0 is the distance between the initial segment and the axis of rotation, R is the rotor radius, and r is the distance between the segment and the axis of rotation.

[0030] The main section comprises the middle portion extending from the root of the blades, followed by the tip. The tip begins, for example, at a distance from the axis of rotation equal to 80% of the rotor radius R. Thus, the reduction in load generated by the blades is primarily based on the chord, torsion, and offset obtained in the middle portion of the main section, while hovering performance is achieved through the shape of the tip of the main section and the location of the highest speed (especially during hovering flight).

[0031] In particular, the twist angle is relatively refined in the middle section and leaves more degrees of freedom at the tip to optimize the aircraft's performance in hovering flight, which is adversely affected by this twisting choice in the middle section. The string law also allows for maximizing the area of ​​the blades that is advantageous in forward flight, while reducing the chord in the inverse circle, which is naturally wider than that of a conventional helicopter at high speeds.

[0032] The synergistic effect of these features makes it possible to obtain a blade that can first generate reasonable forces on the pitch control system or even on the frequency adapter or drag damper, and secondly achieve acceptable performance at low speeds.

[0033] For example, this type of blade allows for performance essentially equivalent to conventional blades at low speeds. The hovering quality factor obtained using this blade is close to that of conventional blades for rotor thrust corresponding to the flight envelope. Furthermore, the force generated by this blade at very high speeds (e.g., 220 knots) is similar to the force generated by conventional blades at high speeds (i.e., 150 knots).

[0034] The blades according to the invention may also include one or more of the following features.

[0035] Therefore, the blade root may include a profile created according to the teachings of document EP3501979A1.

[0036] According to one possibility compatible with the foregoing possibilities, the main section may include a first profile from the initial section to an inner section at a distance between 30% and 35% of the rotor radius R from the axis of rotation, a second profile from the inner section to an intermediate section at a distance equal to, for example, 70% of the rotor radius R from the axis of rotation, a second profile from the intermediate section to a transition section at a distance equal to, for example, 90% of the rotor radius R from the axis of rotation, a third profile from the transition section to the final section, and a fourth profile in the final section.

[0037] The first profile, second profile, third profile, and fourth profile are different, or may even be “OA” profiles known to those skilled in the art. For example, the first profile may be in the form of OA415 profile, the second profile may be in the form of OA312 profile, the third profile may be in the form of OA309 profile, and the fourth profile may be in the form of OA407 profile.

[0038] According to a possibility compatible with the foregoing possibilities, regardless of the implementation method, the blade may include all of the following features:

[0039] The chord length of the segment increases from the initial segment to the maximum chord length reached at an average increase rate of 4.47%.

[0040] The initial section has a relative thickness of 0.45. The main part has a relative thickness that starts from the initial section, which linearly decreases from a value of 0.15 to 0.12 in the inner section, and then remains constant until the middle section, which is 70% of the rotor radius R.

[0041] The twist angle increases according to the law of convexity at the blade root, and the first gradient is equal to -10° divided by the rotor radius R; and

[0042] The quarter chord is located on the pitch axis at the starting section, and its dimensionless offset distance relative to the rotor radius R is equal to +0.041% of the rotor radius in the segment extending from the initial section to the break section. The quarter chord of this segment is located between the pitch axis and the leading edge of the segment.

[0043] The synergistic effect of the above features makes it possible to obtain a blade that can generate reasonable force at high speeds in the pitch control system or even in the frequency adapter or drag damper.

[0044] In addition, the blades may include the following features:

[0045] The initial section is set at a distance of 24% of the rotor radius R from the axis of rotation; and

[0046] The dimensionless average aerodynamic chord relative to the rotor radius R is equal to 0.0651.

[0047] In particular, the following variants have different tips.

[0048] According to the first variant:

[0049] The dimensionless maximum chord relative to the rotor radius R is equal to 0.0746, and the first distance is equal to 78.2% of the rotor radius R;

[0050] The second distance is equal to 87% of the rotor radius R. The chord of the main section decreases from -2.85% to -18.16% in the second section. The chord decreases along the curve from the second section. The curve has an inflection point in the inflection point section located at the third distance, which is equal to 94.5% of the rotor radius R from the axis of rotation. The slope of the curve at the inflection point is along the horizontal axis.

[0051] The relative thickness of the main section decreases linearly away from the middle section until it reaches a relative thickness of 0.07 in the final section;

[0052] The second gradient is equal to -7.3° divided by the rotor radius R, the third gradient is equal to -20° divided by the rotor radius R, and the distance of the fourth section from the axis of rotation is equal to 88% of the rotor radius R; and

[0053] The distance from the fractured section to the axis of rotation is equal to 85% of the rotor radius R. The dimensionless offset distance relative to the rotor radius R is equal to 0.051 in the final section, where the quarter chord point is located between the pitch axis and the trailing edge. The dimensionless offset distance relative to the rotor radius R varies from the fractured section to the final section according to the hyperbolic tangent law.

[0054] For example, the hyperbolic tangent law is defined by the following equation:

[0055]

[0056] Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant section from the axis of rotation. tip "Equals -0.051 multiplied by rotor radius R, "YAC0" equals 0.041 multiplied by rotor radius R, "r" D / R” equals 0.85, and “k” equals 23.1.

[0057] When hovering, this first variant can achieve a quality factor substantially equal to or even better than that of a conventional blade with a constant chord for the rotor thrust corresponding to the operating flight envelope. Furthermore, a decrease in the quality factor may occur for higher thrust compared to the conventional blade and to operational requirements. This later decrease in the quality factor offers the possibility of expanding operational requirements, for example, as part of an upgrade to the aircraft equipped with this blade. For instance, this later decrease in the quality factor can increase the weight of the aircraft.

[0058] According to the second variant:

[0059] The dimensionless maximum chord relative to the rotor radius R is equal to 0.0746, and the first distance is equal to 78.2% of the rotor radius R;

[0060] The second distance is equal to 87% of the rotor radius R. In the second section S2, the decrease occurs from -2.85% to -18.16%. The chord decreases from the second section along the curve with an inflection point in the third distance, which is equal to 94.5% of the rotor radius R from the axis of rotation. The slope of the curve at the inflection point is along the tilt axis.

[0061] The relative thickness of the main section decreases linearly with distance from the middle section to a thickness equal to 0.08 of the section at a distance from the axis of rotation equal to 90% of the rotor radius R, up to and including the final section;

[0062] The second gradient is equal to -7.3° divided by the rotor radius R, the third gradient is equal to -20° divided by the rotor radius R, and the distance of the fourth section from the axis of rotation is equal to 88% of the rotor radius R; and

[0063] The distance from the fractured section to the axis of rotation is equal to 91.5% of the rotor radius R. The dimensionless offset distance relative to the rotor radius R is equal to 0.03384 in the final section, where the quarter chord point is located between the pitch axis and the trailing edge. The dimensionless offset distance relative to the rotor radius R varies from the fractured section to the final section according to the hyperbolic tangent law.

[0064] Alternatively, the hyperbolic tangent law is defined by the following equation:

[0065]

[0066] Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant section from the axis of rotation. tip "Equals -0.03384 multiplied by rotor radius R, "YAC0" equals 0.041 multiplied by rotor radius R, "r" D / R” equals 0.85, and “k” equals 31.

[0067] When hovering, the second variant can achieve a quality factor that is essentially the same as that of the first variant, but the quality factor decreases earlier.

[0068] According to the third variant:

[0069] The dimensionless maximum chord relative to the rotor radius is 0.0746, and the first distance is 78.2% of the rotor radius R.

[0070] The second distance is equal to 94.5% of the rotor radius R;

[0071] The relative thickness of the main section decreases linearly with distance from the middle section to a relative thickness of 0.07 in the final section;

[0072] The second gradient is equal to -10° divided by the rotor radius R, the third gradient is equal to -20° divided by the rotor radius R, and the distance of the fourth section from the axis of rotation is equal to 88% of the rotor radius R; and

[0073] The distance from the fractured section to the rotation axis is equal to 85% of the rotor radius R. The dimensionless offset distance relative to the rotor radius R is equal to 0.051 in the final section, where the quarter chord point is located between the pitch axis and the trailing edge. The dimensionless offset distance relative to the rotor radius R varies from the fractured section to the final section according to a predetermined law.

[0074] Alternatively, the predetermined law is defined by the following equation:

[0075]

[0076] Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant section from the axis of rotation. tip "Equals -0.051 multiplied by rotor radius R, "YAC0" equals 0.041 multiplied by rotor radius R, "r" D / R” equals 0.85, “k” equals 23.1, “d” equals -0.29% of rotor radius R, “u” equals 92.8% of rotor radius R, and “sig” equals 3.7% of rotor radius R.

[0077] When hovering, the third variant has a reduced quality factor compared to the first and second variants, which correspond to the rotor thrust of the operating flight envelope, but maintains the quality factor for thrust greater than the operating requirements. This can provide the possibility of expanding the operating requirements, for example, in the context of upgrading the aircraft equipped with the blade.

[0078] According to the fourth variant:

[0079] The dimensionless maximum chord relative to the rotor radius R is 0.0746, and the first distance is 78.2% of the rotor radius R.

[0080] The second distance is equal to 94.5% of the rotor radius R;

[0081] The relative thickness of the main section decreases linearly with distance from the middle section to a relative thickness of 0.07 in the final section;

[0082] The second gradient is equal to -10° divided by the rotor radius R, the third gradient is equal to -20° divided by the rotor radius R, and the distance of the fourth section from the axis of rotation is equal to 88% of the rotor radius R; and

[0083] The distance from the fractured section to the rotation axis is equal to 94% of the rotor radius R. The dimensionless offset distance relative to the rotor radius R is equal to 0.03384 in the final section, where the quarter chord point is located between the pitch axis and the trailing edge. The dimensionless offset distance relative to the rotor radius R changes from the fractured section to the final section according to a predetermined law.

[0084] Alternatively, the predetermined law is defined by the following equation:

[0085]

[0086] Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant section from the axis of rotation. tip "Equals -0.03384 multiplied by rotor radius R, "YAC0" equals 0.041 multiplied by rotor radius R, "r" D " / R" equals 0.85, "k" equals 32, "d" equals 0.15% of the rotor radius R, "u" equals 95.5% of the rotor radius R, and "sig" equals 1.3% of the rotor radius R.

[0087] The fourth variant has similar performance to the third variant.

[0088] Regardless of the implementation method, the main part may include an intermediate portion starting from the root of the blade, followed by a tip with a zero dihedral angle relative to the intermediate portion.

[0089] Alternatively, dihedral blades are possible. Blades with dihedral blades can achieve more advantageous performance during hovering flight.

[0090] The present invention also relates to a rotor having a hub that is rotatably movable about a rotation axis, the rotor including at least one blade according to the invention attached to the hub.

[0091] An aircraft may be equipped with such a rotor. The aircraft may include a pitch control system for controlling the pitch of each blade of the rotor. Attached Figure Description

[0092] The invention and its advantages will become more apparent from the following description, which is given by way of example with reference to the accompanying drawings, wherein:

[0093] Figure 1 This is a partial view of the blades according to the invention arranged on the rotor of an aircraft;

[0094] Figure 2 This is a view of an exemplary blade according to the present invention;

[0095] Figure 3 This is a view illustrating the outline of the blade according to the present invention;

[0096] Figure 4 It is shown Figure 2 A graph showing the dimensionless variation of the chord of a blade segment relative to the rotor radius;

[0097] Figure 5 It is shown Figure 2 A graph showing the variation in the relative thickness of different sections of the blade;

[0098] Figure 6 It is shown Figure 2 A diagram showing the variation of the twist angle of a section of the blade;

[0099] Figure 7 It is shown Figure 2 A graph showing the variation in the offset of a section of the blade;

[0100] Figure 8 This is a view of an exemplary blade according to the present invention;

[0101] Figure 9 It is shown Figure 8 A graph showing the dimensionless variation of the chord of a blade segment relative to the rotor radius;

[0102] Figure 10 It is shown Figure 8 A graph showing the variation in the relative thickness of different sections of the blade;

[0103] Figure 11 It is shown Figure 8 A graph showing the variation in the offset of a section of the blade;

[0104] Figure 12 This is a view of an exemplary blade according to the present invention;

[0105] Figure 13 It is shown Figure 13A graph showing the dimensionless variation of the chord of a blade segment relative to the rotor radius;

[0106] Figure 14 It is shown Figure 13 A diagram showing the variation of the twist angle of a section of the blade;

[0107] Figure 15 It is shown Figure 13 A graph showing the variation in the offset of a section of the blade;

[0108] Figure 16 This is a view of an exemplary blade according to the present invention; and

[0109] Figure 17 It is shown Figure 16 A graph showing the variation in the offset of a section of the blade. Detailed Implementation

[0110] Elements present in more than one figure are given the same reference numerals in each of them.

[0111] Figure 1 The blade 20 according to the invention is shown schematically. Each blade 20 is a component of the rotor 10 of the aircraft 1.

[0112] Each blade 20 can be carried, for example, via a sleeve 12 through a hub 11. As shown in the illustration, the blade 20 is fixed to the sleeve 12, which is hinged by a laminated stop 9 at the hub 11. The hub 11 is constrained to rotate together with the rotor shaft 13, which causes the hub 11 and blade 20 to rotate about the axis of rotation AX.

[0113] Furthermore, each blade 20 is also capable of rotational movement, particularly around its own pitch axis AXPAS or even around the drag axis. The pitch axis AXPAS extends essentially in a vertical plane radial to the rotation axis AX. To control the blade pitch, the aircraft 1 includes a pitch control system 2.

[0114] The pitch control system 2 includes, for example, several actuators 6 hinged to a lower plate 5 of a set of conventional swashplates. Each blade 20 is then hinged to a pitch control rod 3, which is also hinged to an upper plate 4 of the same set of swashplates. An upper scissor link 7 can be hinged to the rotor shaft 13 and the upper plate 4, and a lower scissor link 8 is hinged to the lower plate 5 and a fixed support for the aircraft 1.

[0115] The aircraft 1 may also include a frequency adapter or a drag damper. Each frequency adapter or drag damper is hinged to the blade 20 and the adjacent blade or hub 11.

[0116] The blade 20 according to the invention can limit the forces applied to the pitch control system 2 and the frequency adapter or drag damper at very high forward speeds of the aircraft 1, while giving the aircraft normal performance at low speeds.

[0117] Figures 2 to 17 An example of the blade 20 according to the present invention is shown.

[0118] Without considering the implementation method and referring to, for example, Figure 2 The blade 20 extends along the pitch axis AXPAS and moves from the first end 51 to the second end 52 away from the rotation axis AX. The distance that separates the second end 52 from the rotation axis AX is called the rotor radius R.

[0119] Specifically, the blade 20 includes a blade body 25, which continuously has a blade root 30 and a subsequent main portion 40 along the pitch axis AXPAS and away from the rotation axis AX. In the given example, the blade root 30 includes a first end portion 51 and can be attached to the sleeve 11. On the other hand, the main portion 40 includes a second end portion 52. The main portion 40 can be decomposed into an intermediate portion 41 extending from a tip 42, extending from the blade root 30 to a distance from the rotation axis AX equal to 80% of the rotor radius R. The tip 42 may have a zero dihedral angle relative to the intermediate portion 41.

[0120] The blade body 25 consists of a series of sections S that are substantially perpendicular to the pitch axis AXPAS. If necessary, the reference numeral S indicates any section, and the reference numerals SO, SF, SREF, SINI, SINT, S1, S2, S3, S4, S5, S6, SINF, and SRUPT indicate special sections.

[0121] Therefore, the blade root 30 extends from the initial section SO forming the first end 51 to a first section called the initial section SINI in the main portion 40. According to the given example, the initial section SINI can be located at a distance from the rotation axis AX between 20% and 30% of the rotor radius R, and advantageously equal to 24% of the rotor radius R. The rotor radius R can be a predetermined characteristic of the blade according to the invention, with only the blade root having a length that varies from one rotor to another.

[0122] The main section 40 extends from the initial section SINI to the final section SF forming the second end 52. The main section 40 also includes a reference section SREF located at a distance from the rotation axis AX equal to 70% of the rotor radius R.

[0123] Each segment S has an aerodynamic profile and extends along a transverse axis from the leading edge BA to the trailing edge BF, with the leading edge BA and the trailing edge BF separated by a certain distance to form a chord C. Each segment S also has:

[0124] The thickness is equal to the relative thickness of its maximum thickness T and the quotient of its chord C;

[0125] The geometric twist angle relative to the reference segment SREF; and

[0126] The offset distance YAC between the variable pitch axis AXPAS and the quarter chord.

[0127] Furthermore, each section S of the blade root 30 may have a thick aerodynamic profile, and each section S of the main portion 40 may have a thin aerodynamic profile. This thin aerodynamic profile has a relative thickness of less than 0.15.

[0128] Figure 3 An example of an aerodynamic profile is shown.

[0129] The blade root 30 may have a circular profile. For example, the blade root 30 may have a first profile P1 from the initial segment SO to segment SO1, and then a second profile P2 with a relatively small thickness, up to but not including the initial segment SINI. For example, the blade root may have a profile created according to the teachings of document EP3501979A1.

[0130] The main section 40, from the initial section SINI (inclusive) to the rotation axis AX at a distance that may be between 30% and 35% (inclusive) of the rotor radius R, the inner section S5 (exclusive) may have a first profile, for example, of type OA415. The main section 40 then, from the inner section S5 (inclusive) to the rotation axis AX at a distance that may be, for example, equal to 70% of the rotor radius R, the intermediate section SINT (exclusive) may have a second profile, for example, of type OA312. The main section 40 then, from the intermediate section SINT (inclusive) to the rotation axis AX at a distance that may be equal to 90% of the rotor radius R, the transition section S6 (exclusive) may have a second profile, then from the transition section S6 (inclusive) to the final section SF (exclusive) has a third profile, for example, of type OA309, and in the final section SF has a fourth profile, for example, of type OA407.

[0131] in addition, Figure 4 The diagram shows the dimensionless chord C of segment S with respect to rotor radius R on the y-axis and the dimensionless distance r relative to rotor radius R that separates each segment of the rotation axis AX on the x-axis. The diagram thus illustrates the laws relating to the variation of the chord of blade 20.

[0132] Regardless of the implementation method, the chord C of segment S increases from the initial segment SO to the maximum chord Cmax. This chord Cmax is reached in the first segment S1, located at a first distance between 75% and 80% (inclusive) of the rotor radius R from the rotation axis AX. Then, the chord C decreases along a curve that first decreases slowly to the second segment S2, and then decreases rapidly. The expressions "slowly decreasing" and "rapidly decreasing" mean that the arithmetic mean of the gradients between two adjacent segments S between the first segment S1 and the second segment S2 is less than the arithmetic mean of the gradients between two adjacent segments S between the second segment S2 and the final segment SF.

[0133] The second section S2 is located at a second distance from the rotation axis AX, between 80% and 95% (inclusive) of the rotor radius R.

[0134] It should be noted that the dimensionless average aerodynamic chord Caero relative to the rotor radius R is between 0.05 and 0.08 (inclusive). Furthermore, the dimensionless maximum chord Cmax relative to the rotor radius R is greater than the dimensionless average aerodynamic chord Caero relative to the rotor radius R, and is between 0.06 and 0.1 (inclusive).

[0135] Specifically, all the examples shown have a dimensionless mean aerodynamic chord Caero with respect to rotor radius R equal to 0.0651. According to another aspect, the chord of segment S can be increased from the chord corresponding to the initial segment SO to the maximum chord Cmax at an average rate of 4.47%.

[0136] Figure 2 The implementation itself has a dimensionless maximum chord Cmax of 0.0746 relative to the rotor radius R, with the first distance equal to 78.2% of the rotor radius R. Furthermore, the second distance is equal to 87% of the rotor radius R. Additionally, the decrease in chord of segment S of the main portion 40 changes from -2.85% to -18.16% at the second segment S2. Then, the law relating to the change in chord has an inflection point PINF in the inflection point segment SINF located at a third distance from the rotation axis AX equal to 94.5% of the rotor radius R, with the slope of the curve at the inflection point along the horizontal axis AXH.

[0137] also, Figure 5 The presented graph shows the relative thickness T / C of segment S on the y-axis and the dimensionless distance r relative to the rotor radius R that separates each segment S from the axis of rotation AX on the x-axis. The graph thus illustrates the laws relating to the variation in the relative thickness of the segments of blade 20.

[0138] Regardless of the implementation method, the relative thickness T / C of segment S decreases from the initial segment S0 to the initial segment SINI, where the relative thickness T / C is between 0.25 and 0.70 (inclusive) up to between 0.12 and 0.15 (inclusive). The main section 40 has a constant relative thickness T / C from the initial segment SINI to the intermediate segment SINT, or decreases according to the given example and then remains constant. Furthermore, the relative thickness T / C of the main section 40 decreases further away from the intermediate segment SINT to a relative thickness T / C between 0.07 and 0.08 (inclusive) in the final segment, where the distance of the intermediate segment SINT from the rotation axis AX is between 65% and 85% (inclusive) of the rotor radius R.

[0139] Specifically, all the examples shown have an initial section SINI with a relative thickness T / C of 0.45. The main portion 40 of these examples linearly decreases the relative thickness T / C from the initial section SINI from a value of 0.15 to a value of 0.12 reached in the inner section S5, located between 30% and 35% (inclusive) of the rotor radius R, and then remains constant in relative thickness T / C until the intermediate section SINT. The intermediate section SINT is located at a distance from the rotation axis AX equal to 70% of the rotor radius R.

[0140] Figure 2 The embodiment shown has a main portion 40 with a relative thickness T / C that decreases linearly away from the intermediate segment SINT until it reaches a relative thickness of 0.07 in the final segment SF.

[0141] also, Figure 6 The presented graph shows the geometric twist angle TETAG of segment S on the y-axis and the dimensionless distance r relative to the rotor radius R that separates each segment S from the axis of rotation AX on the x-axis. The graph thus illustrates the laws relating to the variation in the torsion of blade 20.

[0142] Regardless of the implementation method, the geometric twist angle of segment S increases continuously from the minimum negative angle TETAGMIN to the maximum positive angle TETAGMAX from the initial segment SO to the initial segment SINI. Then, the geometric twist angle of the segment decreases with a first gradient GRAD1 between -9° divided by the rotor radius R and -13° divided by the rotor radius R (inclusive) to the third segment S3 located at a distance between 70% and 80% (inclusive) of the rotor radius R from the rotation axis AX. Then, it reaches the fourth segment S4 located at a distance between 88% and 90% (inclusive) of the rotor radius R from the rotation axis AX with a second gradient GRAD2 equal to or greater than the first gradient GRAD1. Finally, it reaches the final segment SF with a third gradient GRAD3 less than the first gradient GRAD1.

[0143] In particular, all the examples shown have a twist angle increased according to the convexity law at the blade root 30, and a first gradient GRAD1 equal to -10° divided by the rotor radius R.

[0144] Figure 2 The implementation has a second gradient GRAD2 equal to -7.3° divided by the rotor radius R, a third gradient GRAD3 equal to -20° divided by the rotor radius R, and a fourth segment S4 at a distance from the rotation axis AX equal to 88% of the rotor radius R.

[0145] also, Figure 7 The graph shows the dimensionless offset distance of segment S relative to the rotor radius R on the y-axis and the dimensionless distance r relative to the rotor radius R that separates each segment S from the rotation axis AX on the x-axis. The graph thus illustrates the laws relating to the variation in the offset of segment S of blade 20.

[0146] Regardless of the implementation method, in each segment S, the offset distance separates the pitch axis AXPAS from the point Pt located at one-quarter of the chord C of that segment S, starting from the leading edge BA. Point Pt in segment S forms a line referred to by those skilled in the art as a quarter-chord.

[0147] The offset distance decreases from the initial segment SO to the initial segment SINI, then remains constant from the initial segment SINI to the rupture segment SRUPT in the main section 40, and finally decreases to the final segment. The distance between the rupture segment SRUPT and the rotation axis AX is between 80% and 95% (inclusive) of the rotor radius R.

[0148] Specifically, all the examples shown have an offset distance at the starting segment SO, which is equal to one-quarter of the chord of that starting segment SO. In other words, the pitch axis and the quarter-chord line coincide in the starting segment SO.

[0149] Furthermore, the dimensionless offset distance relative to the rotor radius R is equal to +0.041% of the rotor radius R in the segment extending from the initial segment SINI to the rupture segment SRUPT, with the quarter chord of this segment between the pitch axis AXPAS and the leading edge BA of that segment.

[0150] Figure 7The implementation includes a fracture section SRUPT located at a distance from the rotation axis AX equal to 85% of the rotor radius R. The dimensionless offset distance relative to the rotor radius R is also equal to 0.051 in the final section SF, where the quarter-chord point lies between the pitch axis AXPAS and the trailing edge BF. The dimensionless offset distance relative to the rotor radius R varies from the fracture section SRUPT to the final section SF according to the hyperbolic tangent law. This hyperbolic tangent law can be expressed in the following form:

[0151]

[0152] Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant section from the axis of rotation. tip "Equals -0.051 multiplied by rotor radius R, "YAC0" equals 0.041 multiplied by rotor radius R, "r" D / R” equals 0.85, and “k” equals 23.1.

[0153] Figures 8 to 17 The implementation method has the same Figures 2 to 7 Different implementation methods are available at the tip 42.

[0154] Figures 8 to 11 The first variant V1 is shown, which is related to Figure 2 The implementation differs in its laws governing the variation of chord C, relative thickness, and offset at the level of the tip 42. The torsional law differs from that previously described. Figures 2 to 7 The laws governing the implementation methods are the same.

[0155] for Figure 2 Implementation methods and references Figure 9 For this first variant V1, the dimensionless maximum chord Cmax relative to the rotor radius R is equal to 0.0746, and the first distance is equal to 78.2% of the rotor radius R. Additionally, the second distance separating the second segment S2 from the rotation axis AX is equal to 87% of the rotor radius R. It should be noted that a decrease from -2.85% to -18.16% occurs at the second segment S2. The chord C decreases from the second segment S2 along a curve with an inflection point PINF in the inflection segment SINF located at a third distance from the rotation axis AX equal to 94.5% of the rotor radius R, but the slope of the curve at the inflection point is along the tilted axis AXO instead of... Figure 2 The horizontal axis of the blade.

[0156] refer to Figure 10The relative thickness T / C of the tip 42 of the main section 40 decreases linearly with distance from the intermediate section SINT to a thickness of 0.08 at the section located at a distance equal to 90% of the rotor radius R from the axis of rotation, and then remains constant until the final section SF is included.

[0157] refer to Figure 11 The distance from the fractured section SRUPT to the rotation axis AX is equal to 91.5% of the rotor radius R. The dimensionless offset distance relative to the rotor radius R is equal to 0.03384 in the final section SF, where the quarter-chord point lies between the pitch axis AXPAS and the trailing edge BF. The dimensionless offset distance relative to the rotor radius R varies from the fractured section SRUPT to the final section SF according to the hyperbolic tangent law. This hyperbolic tangent law can be expressed in the following form:

[0158]

[0159] Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant section from the axis of rotation. tip "Equals -0.03384 multiplied by rotor radius R, "YAC0" equals 0.041 multiplied by rotor radius R, "r" D / R” equals 0.85, and “k” equals 31.

[0160] Figures 12 to 15 The second variant V2 is shown, which is related to Figures 2 to 7 The implementation differs in its laws governing the variation of chord, torsion, and offset at the tip 42. The law governing the variation of relative thickness differs from that previously described. Figures 2 to 7 The laws governing the implementation methods are the same.

[0161] refer to Figure 13 ,for Figures 2 to 11 The blade has a dimensionless maximum chord Cmax relative to the rotor radius R, which is equal to 0.0746. The first distance is equal to 78.2% of the rotor radius R. On the other hand, the second distance is equal to 94.5% of the rotor radius R.

[0162] refer to Figure 14 The second variant V2 and Figures 2 to 7 The implementation differs in that it has a second gradient GRAD2 equal to -10° divided by the rotor radius R, a third gradient GRAD3 equal to -20° divided by the rotor radius R, and the fourth segment S4 is also at a distance from the rotation axis AX equal to 88% of the rotor radius R.

[0163] refer to Figure 15 ,for Figures 2 to 7The blade 20, with the ruptured section SRUPT at a distance from the rotation axis AX equal to 85% of the rotor radius R, has a dimensionless offset distance relative to R equal to 0.051 in the final section SF, where the quarter-chord point lies between the pitch axis AXPAS and the trailing edge BF. On the other hand, the dimensionless offset distance relative to the rotor radius R varies according to a predetermined law from the ruptured section SRUPT to the final section SF. This predetermined law is:

[0164]

[0165] Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant section from the axis of rotation. tip "Equals -0.051 multiplied by rotor radius R, "YAC0" equals 0.041 multiplied by rotor radius R, "r" D / R” equals 0.85, “k” equals 23.1, “d” equals -0.29% of rotor radius R, “u” equals 92.8% of rotor radius R, and “sig” equals 3.7% of rotor radius R.

[0166] Figures 16 to 17 The third variant V3 is shown, which is related to Figures 2 to 7 The only difference in the implementation method is its offset law at the tip 42.

[0167] Therefore, the blade 20 includes a dimensionless maximum chord Cmax equal to 0.0746 relative to the rotor radius R, a first distance equal to 78.2% of the rotor radius R, and a second distance equal to 94.5% of the rotor radius R.

[0168] The relative thickness T / C of the main section 40 decreases linearly with distance from the intermediate section SINT until it reaches a relative thickness T / C of 0.07 in the final section SF.

[0169] Furthermore, the second gradient GRAD2 is equal to -10° divided by the rotor radius R, the third gradient GRAD3 is equal to -20° divided by the rotor radius R, and the distance between the fourth segment S4 and the rotation axis AX is equal to 88% of the rotor radius R.

[0170] On the other hand, reference Figure 17 The distance between the fractured section SRUPT and the rotation axis AX is now equal to 94% of the rotor radius R. The dimensionless offset distance relative to the rotor radius R is equal to 0.03384 in the final section SF, where the quarter-chord point lies between the pitch axis AXPAS and the trailing edge BF. Finally, the dimensionless offset distance relative to the rotor radius R varies from the fractured section SRUPT to the final section SF according to the following predetermined law:

[0171]

[0172] Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant section from the axis of rotation. tip "Equals -0.03384 multiplied by rotor radius R, "YAC0" equals 0.041 multiplied by rotor radius R, "r" D " / R" equals 0.85, "k" equals 32, "d" equals 0.15% of the rotor radius R, "u" equals 95.5% of the rotor radius R, and "sig" equals 1.3% of the rotor radius R.

[0173] Of course, the invention can be subject to many variations in its implementation. Although several embodiments have been described above, it should be readily understood that not all possible embodiments can be exhaustively determined. Equivalent devices can, of course, be substituted for any of the described devices without departing from the scope of the invention and the claims, which define the invention.

Claims

1. A blade (20) for a rotor (10) of an aircraft (1), the blade (20) being rotatable about a rotation axis (AX) and a pitch axis (AXPAS) of the rotor (10), the blade (20) extending along the pitch axis (AXPAS) from a first end (51) to a second end (52), the blade (20) comprising a blade body (25) having a blade root (30) along the pitch axis (AXPAS) and then a main portion formed by a series of segments (S) substantially perpendicular to the pitch axis (AXPAS), the blade root (30) being provided with forming the first end (51) (52). 1) The initial section (SO), the main part (40) extends along the pitch axis (AXPAS) from the initial section (SINI) to the final section (SF) forming the second end (52), the final section (SF) being located at a distance equal to a predetermined rotor radius (R) of the rotation axis (AX), each section extending along the transverse axis from the leading edge (BA) to the trailing edge (BF), the leading edge (BA) and the trailing edge (BF) being separated by the maximum distance constituting the chord (C), each section of the blade body (25) having a geometric twist angle relative to a reference section (SREF) located at a distance equal to 70% of the rotor radius (R) from the rotation axis, in: The initial section (SINI) is located at a distance between 20% and 30% of the rotor radius (R) from the axis of rotation (AX); The chord (C) of the segment increases from the initial segment (S0) to the maximum chord (Cmax) reached in the first segment (S1), and then decreases according to the law of first slowly decreasing to the second segment (S2) and then rapidly decreasing. The first segment (S1) is located at a first distance from the axis of rotation (AX) between 75% and 80% of the rotor radius (R), and the second segment (S2) is located at a second distance from the axis of rotation (AX) between 80% and 95% of the rotor radius (R). The dimensionless average aerodynamic chord (Caero) relative to the rotor radius (R) is between 0.05 and 0.08, and the dimensionless maximum chord (Cmax) relative to the rotor radius (R) is greater than the dimensionless average aerodynamic chord (Caero) relative to the rotor radius (R) and is between 0.06 and 0.

1. The relative thickness (T / C) of the segment decreases from the starting segment (SO) to the initial segment (SINI) from a relative thickness (T / C) between 0.25 and 0.70 to a relative thickness (T / C) between 0.12 and 0.

15. The main part (40) has a constant relative thickness (T / C) or decreases but then remains constant from the initial segment (SINI) to the intermediate segment (SINT). The relative thickness (T / C) of the main part (40) decreases as it moves away from the intermediate segment (SINT) to a relative thickness (T / C) between 0.07 and 0.08 in the final segment. The distance of the intermediate segment (SINT) from the axis of rotation (AX) is between 65% and 85% of the rotor radius (R). The geometric twist angle increases continuously from the minimum negative angle (TETAGMIN) to the maximum positive angle (TETAGMAX) from the starting segment (SO) to the initial segment (SINI), then decreases with a first gradient (GRAD1) between -9° divided by the rotor radius and -13° divided by the rotor radius (R), to a third segment (S3) located at a distance from the rotation axis (AX) equal to or greater than 70% and 80% of the rotor radius (R), then reaches a fourth segment (S4) located at a distance from the rotation axis (AX) equal to or greater than 88% and 90% of the rotor radius (R), and finally reaches the final segment (SF) with a third gradient (GRAD3) less than the first gradient (GRAD1). The offset distance (YAC) separating the pitch axis (AXPAS) from the quarter chord of each segment decreases from the starting segment (SO) to the initial segment (SINI), the offset distance being constant from the initial segment (SINI) to the rupture segment (SRUPT) in the main section (40) and then decreasing, the distance of the rupture segment (SRUPT) from the rotation axis (AX) being between 80% and 95% of the rotor radius (R).

2. The blade according to claim 1, wherein, The main portion (40) includes: a first profile from the initial section (SINI) to an inner section (S5) located at a distance from the rotation axis (AX) between 30% and 35% of the rotor radius (R); a second profile from the inner section (S5) to the intermediate section (SINT), the second profile extending from the intermediate section (SINT) to the transition section (S6); a third profile from the transition section (S6) to the final section (SF); and a fourth profile within the final section (SF).

3. The blade according to claim 1, wherein, The blade includes the following features: The chord of the segment increases from the initial segment (SO) to the maximum chord (Cmax) reached at an average increase rate of 4.47%. The initial section (SINI) has a relative thickness (T / C) of 0.45, and the main section (40) has a relative thickness (T / C) that decreases linearly from 0.15 to 0.12 at the inner section (S5) and remains constant until the intermediate section (SINT), which is located at a distance from the axis of rotation (AX) equal to 70% of the rotor radius (R). The twist angle increases according to the convexity law at the blade root (30), and the first gradient (GRAD1) is equal to -10° divided by the rotor radius (R); and The quarter chord is located on the pitch axis at the starting section (SO), and the offset distance is equal to +0.041% of the rotor radius (R) in the segment extending from the initial section (SINI) to the rupture section (SRUPT), and the quarter chord of the segment is located between the pitch axis (AXPAS) and the leading edge (BA) of the segment.

4. The blade according to claim 3, wherein, The blade includes the following features: The initial section (SINI) is located at a distance from the axis of rotation (AX) equal to 24% of the rotor radius (R); and The dimensionless average aerodynamic chord (Caero) relative to the rotor radius (R) is equal to 0.0651.

5. The blade according to claim 1, wherein: The dimensionless maximum chord (Cmax) relative to the rotor radius (R) is equal to 0.0746, and the first distance is equal to 78.2% of the rotor radius (R); The second distance is equal to 87% of the rotor radius (R), the chord of the segment of the main part (40) decreases from -2.85% to -18.16% at the second segment (S2), the chord (C) decreases along the curve (C1) from the second segment (S2), the curve (C1) has an inflection point (PINF) in the inflection point segment (SINF) located at a third distance from the axis of rotation (AX) equal to 94.5% of the rotor radius (R), the slope of the curve at the inflection point is along the horizontal axis (AXH); The relative thickness (T / C) of the main section (40) decreases linearly with distance from the intermediate section (SINT) to a relative thickness of 0.07 in the final section (SF); The second gradient (GRAD2) is equal to -7.3° divided by the rotor radius (R), the third gradient (GRAD3) is equal to -20° divided by the rotor radius (R), and the fourth segment (S4) is at a distance from the rotation axis (AX) equal to 88% of the rotor radius (R); and The distance of the fractured section (SRUPT) from the axis of rotation (AX) is equal to 85% of the rotor radius (R), and the dimensionless offset distance relative to the rotor radius (R) is equal to 0.051 in the final section (SF), where the quarter chord point is located between the pitch axis (AXPAS) and the trailing edge (BF), and the dimensionless offset distance relative to the rotor radius (R) varies from the fractured section (SRUPT) to the final section (SF) according to the hyperbolic tangent law.

6. The blade according to claim 5, wherein, The hyperbolic tangent law is defined by the following equation: Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant segment from the rotation axis. tip "YAC0" equals -0.051 multiplied by the rotor radius R, "r" equals 0.041 multiplied by the rotor radius R, "r" D / R” equals 0.85, and “k” equals 23.

1.

7. The blade according to claim 1, wherein: The dimensionless maximum chord (Cmax) relative to the rotor radius is equal to 0.0746, and the first distance is equal to 78.2% of the rotor radius (R); The second distance is equal to 87% of the rotor radius (R), and a decrease from -2.85% to -18.16% occurs in the second segment (S2). The chord (C) decreases from the second segment (S2) along a curve with an inflection point (PINF) in the inflection point segment (SINF) located at a third distance equal to 94.5% of the rotor radius (R) from the axis of rotation (AX). The slope of the curve at the inflection point is along the tilt axis (AXO). The relative thickness (T / C) of the main section (40) decreases linearly with distance from the intermediate section (SINT) to a thickness equal to 0.08 of the section at a distance from the axis of rotation equal to 90% of the rotor radius, up to and including the final section (SF). The second gradient (GRAD2) is equal to -7.3° divided by the rotor radius (R), the third gradient (GRAD3) is equal to -20° divided by the rotor radius (R), and the fourth segment (S4) is at a distance from the rotation axis (AX) equal to 88% of the rotor radius (R); and The distance of the fractured section (SRUPT) from the axis of rotation (AX) is equal to 91.5% of the rotor radius (R), and the dimensionless offset distance relative to the rotor radius (R) is equal to 0.03384 in the final section (SF), whereby... The quarter chord point is located between the pitch axis (AXPAS) and the trailing edge (BF), and the dimensionless offset distance relative to the rotor radius (R) varies from the fractured section (SRUPT) to the final section (SF) according to the hyperbolic tangent law.

8. The blade according to claim 7, wherein, The hyperbolic tangent law is defined by the following equation: Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant segment from the rotation axis. tip "YAC0" equals -0.03384 multiplied by the rotor radius R, "r" equals 0.041 multiplied by the rotor radius R, "r" D / R” equals 0.85, and “k” equals 31.

9. The blade according to claim 1, wherein: The dimensionless maximum chord (Cmax) relative to the rotor radius (R) is equal to 0.0746, and the first distance is equal to 78.2% of the rotor radius (R); The second distance is equal to 94.5% of the rotor radius (R); The relative thickness of the main section (40) decreases linearly with distance from the intermediate section (SINT) to a relative thickness (T / C) of 0.07 in the final section (SF); The second gradient (GRAD2) is equal to -10° divided by the rotor radius (R), the third gradient (GRAD3) is equal to -20° divided by the rotor radius (R), and the fourth segment (S4) is at a distance of 88% of the rotor radius (R) from the axis of rotation (AX); and The distance of the fractured section (SRUPT) from the axis of rotation (AX) is equal to 85% of the rotor radius (R), and the dimensionless offset distance relative to the rotor radius (R) is equal to 0.051 in the final section (SF), wherein the quarter chord point is located between the pitch axis (AXPAS) and the trailing edge (BF), and the dimensionless offset distance relative to the rotor radius (R) varies from the fractured section (SRUPT) to the final section (SF) according to a predetermined law.

10. The blade according to claim 9, wherein, The predetermined law is defined by the following equation: Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant segment from the rotation axis. tip "YAC0" equals -0.051 multiplied by the rotor radius R, "r" equals 0.041 multiplied by the rotor radius R, "r" D / R” equals 0.85, "k" equals 23.1, "d" equals -0.29% of rotor radius R, "u" equals 92.8% of rotor radius R, and "sig" equals 3.7% of rotor radius R.

11. The blade according to claim 1, wherein: The dimensionless maximum chord (Cmax) relative to the rotor radius (R) is equal to 0.0746, and the first distance is equal to 78.2% of the rotor radius (R); The second distance is equal to 94.5% of the rotor radius (R); The relative thickness (T / C) of the main section (40) decreases linearly with distance from the intermediate section (SINT) to a relative thickness (T / C) of 0.07 in the final section (SF); The second gradient (GRAD2) is equal to -10° divided by the rotor radius (R), the third gradient (GRAD3) is equal to -20° divided by the rotor radius (R), and the fourth segment (S4) is at a distance of 88% of the rotor radius (R) from the axis of rotation (AX); and The distance of the fractured section (SRUPT) from the axis of rotation (AX) is equal to 94% of the rotor radius (R), and the dimensionless offset distance relative to the rotor radius (R) is equal to 0.03384 in the final section (SF), whereby... The quarter chord point is located between the pitch axis (AXPAS) and the trailing edge (BF), and the dimensionless offset distance relative to the rotor radius (R) varies from the break section (SRUPT) to the final section (SF) according to a predetermined law.

12. The blade according to claim 11, wherein, The predetermined law is defined by the following equation: Where "YAC" is the offset distance, "R" is the rotor radius, and "r" is the radius separating the relevant segment from the rotation axis. tip "YAC0" equals -0.03384 multiplied by the rotor radius R, "r" equals 0.041 multiplied by the rotor radius R, "r" D " / R" equals 0.85, "k" equals 32, "d" equals 0.15% of the rotor radius R, "u" equals 95.5% of the rotor radius R, and "sig" equals 1.3% of the rotor radius R.

13. The blade according to claim 1, wherein, The main part (40) begins at the blade root (30) and includes a middle part (41) followed by a tip (42) having a zero dihedral angle relative to the middle part (41).

14. A rotor (10) having a hub (11) rotatable about a rotation axis (AX), the rotor (10) including at least one blade (20) attached to the hub (11), wherein the blade (20) is the blade according to claim 1.

15. An aircraft (1) equipped with a rotor (10), wherein, The rotor is the rotor according to claim 14.

Images