Aircraft rotor blade of shape adapted for acoustic improvement during an approach flight and for improving performance in forward flight

Abstract

A blade of a rotor for a rotary wing aircraft. The blade presents relationships for variation in the sweep and in the chord of the profiles of sections of the blade, in particular in order to improve the twisting stiffness and the bending stiffness of the blade. The blade is then double-tapered and it presents three sweeps, making it possible firstly to improve the aerodynamic performance of the blade in forward flight, and secondly to reduce the noise given off by the blade, in particular during an approach flight.

Claims

1. A blade for a rotor of a rotary wing aircraft, the blade being for rotating about an axis of rotation (A), the blade extending firstly along a blade axis (B) between a blade start suitable for being connected to a hub of the rotor and a blade tip situated at a free end of the blade, and secondly along a transverse axis (T) perpendicular to the blade axis (B) between a leading edge and a trailing edge, the blade comprising: an airfoil portion situated between the blade start and the blade tip, the airfoil portion being constituted by a succession of airfoil profiles, each airfoil profile being situated in a transverse plane substantially perpendicular to the blade axis (B) and defining a section of the blade, the blade tip being situated at a distance equal to a rotor radius R from the axis of rotation (A), a maximum distance between the leading edge and the trailing edge in the transverse plane constituting a chord c for the airfoil profile of each of the sections of the blade, a mean chord c being a mean value of the chord cover the airfoil portion, a forward first direction being defined from the trailing edge to the leading edge, and a rearward second direction being defined from the leading edge to the trailing edge, the blade presenting a combination of relationships for variation in chord and in sweep, the sweep being the angle between the leading edge and the blade axis (B), the chord increasing between the start of the airfoil portion and a first section S1 situated at a first distance from the axis of rotation (A) lying in the range 0.6R to 0.9R, the chord decreasing beyond the first section S1; and wherein the sweep is directed in the forward first direction of the blade between the start of the airfoil portion and a second section S2 situated at a second distance from the axis of rotation (A) lying between 0.5R and 0.8R, the leading edge forming a forward first sweep angle .sub.1 that is strictly greater than 0 and less than 10 relative to the blade axis (B), the sweep being directed in the forward first direction of the blade between the second section S2 and a third section S3 situated at a third distance from the axis of rotation (A) lying in the range 0.6R to 0.95R, the leading edge forming a forward second sweep angle .sub.2 lying in the range 1 to 15 relative to the blade axis (B), the sweep being directed in the rearward second direction of the blade between the third section S3 and the blade tip, the leading edge forming a backward third sweep angle .sub.3 lying in the range 35 to 15 relative to the blade axis (B), wherein the forward first sweep angle .sub.1 is different from the forward second sweep angle .sub.2, wherein the forward first sweep angle .sub.1, the forward second sweep angle .sub.2, and the backward third sweep angle .sub.3 are constant respectively between the start of the airfoil portion and the second section S2, between the second section S2 and the third section S3, and between the third section S3 and the blade tip.

2. The blade according to claim 1, wherein the forward first sweep angle .sub.1 is strictly less than the forward second sweep angle .sub.2.

3. The blade according to claim 1, wherein the forward first sweep angle .sub.1 is equal to 4, the forward second sweep angle .sub.2 is equal to 8, and the backward third sweep angle .sub.3 is equal to 23.

4. The blade according to claim 1, wherein the blade start is situated at a fourth distance lying in the range 0.05R to 0.3R from the axis of rotation (A) and the start of the airfoil portion is situated at a fifth distance lying in the range 0.1R to 0.4R from the axis of rotation (A), the fifth distance being greater than or equal to the fourth distance, and the chord in the vicinity of the start of the blade lying in the range 0.4c to 0.9c.

5. The blade according to claim 1, wherein the chord varies about the mean chord c by 40% between the start of the airfoil portion and the first section S1.

6. The blade according to claim 1, wherein the chord decreases in non-linear manner beyond a sixth section S6 situated at an eighth distance from the axis of rotation (A) lying in the range 0.9R to 0.95R to the blade tip.

7. The blade according to claim 6, wherein the chord decreases in parabolic manner beyond the sixth section S6.

8. The blade according to claim 1, wherein the blade has a dihedral in the vicinity of the blade tip.

9. The blade according to claim 1, wherein the mean chord c is defined by a radius squared r.sup.2 weighting of the profile of each of the sections of the blade in application of the formula: $\stackrel{\_}{c}=\frac{{}_{R_0}^{R}L\left(r\right).Math.r^2.Math.\mathrm{dr}}{{}_{R_0}^R{r}^2.Math.\mathrm{dr}}$ where L(r) is the length of the local chord of a profile of the blade, the local profile being situated at a radius r from the axis of rotation A, R.sub.0 being the radius of the start of the airfoil portion, and R being the radius of the blade tip.

10. A rotor for a rotary wing aircraft, the rotor having at least two blades according to claim 1.

11. A blade for a rotor of a rotary wing aircraft, the blade rotating about an axis of rotation and extending radially from the axis of rotation along a blade axis to a blade tip a distance equal to a rotor radius R, the blade having a chord dimension perpendicular to the blade axis between a leading edge and a trailing edge, the blade varying in the chord dimension and in sweep, the sweep being the angle between the leading edge and the blade axis, the blade comprising: an increasing chord section extending from a blade start to a first distance from the axis of rotation located between 0.6R to 0.9R along which the chord dimension continuously increases; a decreasing chord section extending from the first distance to the blade tip along which the chord dimension continuously decreases; a first sweep section extending from a start of a blade airfoil shape to a second distance from the axis of rotation located between 0.5R and 0.8R at a forward first sweep angle greater than 0 and less than 10 relative to the blade axis in a forward direction of the blade; a second sweep section extending from the second distance to a third distance from the axis of rotation located between 0.6R and 0.95R at a forward second sweep angle in the range of 1 to 15 relative to the blade axis in the forward direction of the blade, wherein the forward first sweep angle is different from the forward second sweep angle; and a third sweep section extending from the third distance to the blade tip, at a backward third sweep angle .sub.3 in the range 35 to 15 relative to the blade axis in the rearward direction of the blade, wherein the forward first sweep angle, the forward second sweep angle, and the backward third sweep angle are constant respectively.

12. The blade according to claim 11, wherein the forward first sweep angle is less than the forward second sweep angle.

13. The blade according to claim 11, wherein the forward first sweep angle is equal to 4, the forward second sweep angle is equal to 8, and the backward third sweep angle is equal to 23.

14. The blade according to claim 11, wherein the blade start is located in the range 0.05R to 0.3R from the axis of rotation and the start of the airfoil shape is located in the range 0.1R to 0.4R from the axis of rotation.

15. The blade according to claim 11, wherein the chord varies about a mean chord dimension by 40% between the start of the blade and the first distance.

16. The blade according to claim 11, wherein the decreasing chord section has a non-linear section extending in the range 0.9R to 0.95R to the blade tip, along which the chord dimension decreases non-linearly.

17. The blade according to claim 16, wherein the chord dimension decreases parabolically in the non-linear section.

18. The blade according to claim 11, further comprising a dihedral portion in the vicinity of the blade tip.

Description

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

(1) The invention and its advantages appear in greater detail from the context of the following description of embodiments given by way of illustration and with reference to the accompanying figures, in which:

(2) FIGS. 1 and 2 show a blade of the invention;

(3) FIG. 3 shows an aircraft having a rotor made up of such blades;

(4) FIG. 4 is a graph plotting variation in the chord of the profiles of the sections of the blade;

(5) FIG. 5 is a graph plotting variation in the sweep of the blade;

(6) FIG. 6 is a graph plotting variation in the twist of the blade; and

(7) FIG. 7 is a graph plotting variation in the twist gradient of the blade.

DETAILED DESCRIPTION OF THE INVENTION

(8) Elements present in more than one of the figures are given the same references in each of them.

(9) FIGS. 1 and 2 show a blade 1 extending firstly spanwise along a blade axis B between a blade start 2 and a blade tip 9 and secondly along a transverse axis T perpendicular to the blade axis B between a leading edge 6 and a trailing edge 7. The blade 1 has an airfoil portion 4 situated between the blade start 2 and the blade tip 9. The airfoil portion 4 is made up of a succession of airfoil profiles 15, each situated in a transverse plane substantially perpendicular to the blade axis B, each profile defining a section of the blade 1. The blade 1 also has a dihedral 5 at the free end of the blade 1, i.e. in the vicinity of the blade tip 9.

(10) The blade 1 is for forming a rotor 11 of a rotary wing aircraft 10, as shown in FIG. 3. The rotor 11 comprises a hub 12 and five blades 1 that are for rotating about an axis of rotation A of the hub 12. Each blade 1 is connected to the hub 12 at the blade start 2.

(11) The rotor 11 is characterized by a rotor radius R, i.e. the distance between the axis of rotation A and the blade tip 9 along the blade axis B. The chord c of the profiles 15 of each section of the blade 1 corresponds to the maximum distance between the leading edge 6 and the trailing edge 7 of the blade 1 in a transverse plane substantially perpendicular to the blade axis B. A mean chord c is defined as being the mean value of the chords c over the airfoil 4. The blade start 2 is situated at a fourth distance equal to 0.1R from the axis of rotation A and the start 3 of the airfoil 4 of the blade 1 is situated at a fifth distance equal to 0.2R from the axis of rotation A.

(12) The blade 1 of the invention presents a combination of relationships for variation in its sweep and in the chord of its profiles 15 firstly in order to reduce the noise given off by each blade 1 of the rotor 11 during an approach flight, and secondly in order to improve the aerodynamic performance of each blade 1 during forward flight of the aircraft 10.

(13) Furthermore, the blade 1 may also present a combination of relationships for variation in its sweep and in the chord of the profiles 15 of its sections together with a relationship for variation in its twist, firstly in order to reduce the noise given off by each blade 1 of the rotor 11 during an approach flight, and secondly in order to improve the aerodynamic performance of each blade 1 both during hovering flight and forward flight of the aircraft 10.

(14) The relationship for variation in the chord, the sweep, and the twist of the profiles 15 of the sections of the blade 1 are plotted respectively in FIGS. 4 to 6. FIG. 7 shows the twist gradient of the blade 1, i.e. the local derivative of the twist along the span of the blade 1 of the rotor 11 of rotor radius R.

(15) The relationship for variation in the chord of the profiles 15 of the sections of the blade 1 shown in FIG. 4 comprises, plotted along the abscissa axis, the ratio of the positions of the profiles 15 of the sections of the blade 1 along the span of the blade 1 relative to the rotor radius R, and up the ordinate axis, the ratio of the chords c of the profiles 15 of the sections of the blade 1 relative to the mean chord c.

(16) The mean chord c is defined by a radius squared r.sup.2 weighting of each profile 15 of the sections of the blade 1 in application of the following formula:

(17) $\stackrel{\_}{c}=\frac{{}_{R_0}^{R}L\left(r\right).Math.r^2.Math.\mathrm{dr}}{{}_{R_0}^R{r}^2.Math.\mathrm{dr}}$
where L(r) is the length of the local chord of a profile of the blade 1 situated at a radius r from the axis of rotation A, R.sub.0 is the radius of the start 3 of the airfoil 4, and R is the radius of the blade tip 9.

(18) In this relationship for variation in the chord, the chord c of the profile 15 of each section of the blade 1 increases between the start 3 of the airfoil portion 4 and a first section S1 situated at a first distance from the axis of rotation A that is equal to 0.85R. Beyond the first section S1, the chord decreases to the blade tip 9. It can be seen that the section c is less than the mean chord c between the start of the airfoil portion of the blade 1 and a fourth section S4 situated at a sixth distance from the axis of rotation A equal to 0.6R. Furthermore, the chord c varies between the start 3 of the airfoil portion 4 and the first section S1 from 0.8c to 1.2c, which represents variation of 20% about the mean chord c. The chord at the blade start is equal to 0.3c.

(19) Thereafter, the chords of the profiles 15 of the sections of the blade 1 are greater than the mean chord c between the fourth section S4 and a fifth section S5 situated at a seventh distance from the axis of rotation A lying in the range 0.85R to 0.95R. Finally, the chords of the profiles 15 of the sections of the blade 1 are less than the mean chord c beyond the fifth section S5 to the blade tip 9.

(20) In addition, the chord c decreases following a curve that is substantially parabolic beyond a sixth section S6 situated at an eighth distance equal to 0.95R. The end of the blade 1 thus forms a parabolic tip cap 8.

(21) The relationship for variation in the sweep of the blade as shown in FIG. 5 defines three sweeps. The ratio of the position of the profile 15 of each section of the blade 1 along the blade axis B over the rotor radius R is plotted along the abscissa axis, and the sweep angle of each of these profiles 15 is plotted up the ordinate axis.

(22) Thus, the sweep is initially directed towards the front of the blade 1 between the start 3 of the airfoil portion 4 and a second section S2 situated at a second distance from the axis of rotation A equal to 0.67R, the leading edge 6 forming a forward first sweep angle .sub.1 equal to 4 relative to the blade axis B. Thereafter, the sweep is directed towards the front of the blade 1 between the second section S2 and a third section S3 situated at a third distance from the axis of rotation A equal to 0.85R, the leading edge 6 forming a forward second sweep angle .sub.2 equal to 8 relative to the blade axis B. Finally, the sweep is directed towards the rear of the blade 1 between the third section S3 and the blade tip 9, the leading edge 6 forming a backward third sweep angle .sub.3 equal to 23 relative to the blade axis B.

(23) Each of the connections between the first, second, and third sweep angles is preferably made with a connection radius in order to avoid having a sharp angle at any of these connections. These connection radii may for example be of the order of 500 millimeters (mm).

(24) Furthermore, the blade 1 has a downwardly-directed dihedral 5 at its free end. This dihedral 5 begins in the vicinity of the sixth section S6 and terminates at the blade tip 9. The dihedral 5 serves mainly to improve the aerodynamic behavior of the blade 1 in hovering flight by reducing the influence of the vortex generated by the preceding blade.

(25) In addition, a relationship for twist of the profiles 15 may be added to the blade 1 in order to improve the aerodynamic performance of the blade 1 both during hovering flight and during forward flight. This relationship for twist of the blade 1 shown in FIG. 6 is a non-linear relationship corresponding to a polynomial curve. The ratio of the position of each profile 15 of the sections of the blade 1 along the span over the rotor radius R is plotted along the abscissa axis, and the twist angle of the profile 15 of each section of the blade 1 is plotted up the ordinate axis.

(26) The twist gradient is shown in FIG. 7 and comprises, along the abscissa axis, the ratio of the position of the profile 15 of each section of the blade 1 along the span of the blade 1 over the rotor radius R, and, up the ordinate axis, the local derivative of the twist of the profile 15.

(27) Initially, the twist angle varies little between the start 3 of the airfoil portion 4 and a seventh section S7 situated at a ninth distance from the axis of rotation A equal to 0.35R. The variation in the twist angle is less than 2 between the start 3 of the airfoil portion 4 and the seventh section S7. The twist angle increases a little and then decreases along the span, the twist gradient being positive in the vicinity of the start 3 of the airfoil portion 4 and decreasing to become negative in the vicinity of the seventh section S7.

(28) Thereafter, the twist angle decreases between the seventh section S7 and an eighth section S8 situated at a tenth distance from the axis of rotation A equal to 0.48R, the twist gradient decreasing to a first plateau equal to 18/R in the vicinity of the eighth section S8.

(29) Thereafter, the twist angle decreases less between the eighth section S8 and a ninth section S9 situated at an eleventh distance from the axis of rotation A equal to 0.78R, the twist gradient increases up to a second plateau equal to 6/R in the vicinity of the ninth section S9. In particular, the twist angle is equal to 0 for a profile 15 of the blade 1 situated at a distance from the axis of rotation A equal to 0.65R.

(30) The twist angle again decreases more between the ninth section S9 and a tenth section S10 situated at a twelfth distance from the axis of rotation A equal to 0.92R, the twist gradient decreasing to a third plateau equal to 13/R in the vicinity of the tenth section S10.

(31) Finally, the twist angle decreases between the tenth section S10 and the blade tip 9, the twist gradient increasing up to a twist gradient equal to 8/R at the blade tip 9.

(32) Naturally, the present invention may be subjected to numerous variations as to its implementation. Although several implementations are described, it will readily be understood that it is not conceivable to identify exhaustively all possible embodiments. It is naturally possible to envisage replacing any of the means described by equivalent means without going beyond the ambit of the present invention.