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

Abstract

A blade of a rotor for a rotary wing aircraft. The blade presents a combination of relationships for variation in the sweep, the chord, and the twist of the airfoil profiles of the sections of the blade in order, firstly to improve the aerodynamic performance of the blade both in forward flight and in stationary flight, and secondly to reduce the noise given off during approach flight. The blade is double-tapered and presents three sweeps. The twist relationship is substantially constant over a first portion of the blade, and then decreases over the remainder of the blade in linear or in non-linear manner. Suitable variation in the gradient of the twist of the blade makes it possible to improve the aerodynamic performance of the blade in forward flight and in hovering 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 c over 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 its chord and in its twist, the twist being formed by angular variations between the airfoil profiles of the blade, 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 the twist decreasing between a second section S2 situated at a second distance from the axis of rotation (A) lying in the range 0.3R to 0.4R and the blade tip, a first gradient of the twist lying in the range 25/R to 4/R 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 0.4R to 0.6R, a second gradient of the twist lying in the range 25/R to 4/R between the third section S3 and a fourth section S4 situated at a fourth distance from the axis of rotation (A) lying in the range 0.65R to 0.85R, a third gradient of the twist lying in the range 16/R to 4/R between the fourth section S4 and a fifth section S5 situated at a fifth distance from the axis of rotation (A) lying in the range 0.85R to 0.95R, a fourth gradient of the twist lying in the range 16/R to 0/R between the fifth section S5 and the blade tip, wherein the relationship for variation in its chord and in its twist are combined with a relationship for variation in its sweep, the sweep, which is the angle between the leading edge and the blade axis (B) being directed towards the front of the blade between the start of the airfoil portion and a ninth section S9 situated at an eleventh 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 towards the front of the blade between the ninth section S9 and a tenth section S10 situated at a twelfth 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 towards the rear of the blade between the tenth section S10 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).

2. A blade according to claim 1, wherein the twist varies piecewise linearly between adjacent pairs of sections from the second, third, fourth, and fifth sections S2, S3, S4, and S5, and between the fifth section S5 and the blade tip.

3. A blade according to claim 2, wherein the first gradient of the twist is less than the second gradient of the twist, the third gradient of the twist is greater than the third gradient of the twist, and the third gradient of the twist is less than the fourth gradient of the twist.

4. A blade according to claim 1, wherein the twist varies in non-linear manner over the airfoil portion, the first gradient of the twist reaching a first plateau lying in the range 25/R to 15/R in the vicinity of the third section S3, the second gradient of the twist reaching a second plateau lying in the range 14/R to 4/R in the vicinity of the fourth section S4, the third gradient of the twist reaching a third plateau lying in the range 16/R to 6/R in the vicinity of the fifth section S5, and the fourth gradient of the twist lying in the range 10/R to 0/R in the vicinity of the blade tip.

5. A blade according to claim 4, wherein the first plateau is equal to 18/R, the second plateau is equal to 6/R, the third plateau is equal to 13/R, and the fourth plateau of the twist is equal to 8/R at the blade tip.

6. A blade according to claim 1, wherein the variation of the twist is less than or equal to 2 between the start of the airfoil portion and the second section S2.

7. A blade according to claim 1, wherein the second distance is equal to 0.35R, the third distance is equal to 0.48R, the fourth distance is equal to 0.78R, and the fifth distance is equal to 0.92R.

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

9. A 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.

10. A blade according to claim 1, wherein the chord decreases in non-linear manner beyond an eighth section S8 situated at a tenth distance from the axis of rotation (A) lying in the range 0.9R to 0.95R to the blade tip.

11. A blade according to claim 10, wherein the chord decreases in parabolic manner beyond the eighth section S8.

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

13. A blade according to claim 1, wherein the forward first sweep angle .sub.1 is different from the forward second sweep angle .sub.2.

14. A blade according to claim 13, wherein the forward first sweep angle .sub.1 is strictly less than the forward second sweep angle .sub.2.

15. A blade according to claim 1, 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 ninth section S9, between the ninth section S9 and the tenth section S10, and between the tenth section S10 and the blade tip.

16. A 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.

17. A 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.

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

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 sixth 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 seventh 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, its chords, and its twist, firstly in order to reduce the noise given off by each blade of the rotor 11 during an approach flight and secondly for improving the aerodynamic performance of each blade 1, both during hovering flight and during forward flight of the aircraft 10.

(13) The blade 1 may also present a combination solely of the relationships for variation in its chords and its twist in order to improve the aerodynamic performance of the rotor 11, both during hovering flight and during forward flight of the aircraft 10, but without taking account of the acoustic behavior of each blade 1.

(14) The blade 1 may also present a combination of relationships for variation in its sweep and in its chords 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 in forward flight of the aircraft 10. The aerodynamic performance of each blade 1 is then optimized mainly for forward flight of the aircraft 10.

(15) 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.

(16) 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.

(17) 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:

(18) $\stackrel{\_}{c}=\frac{{}_{R_0}^{R}L\left(r\right).Math.r^2.Math.\mathrm{dr}}{{}_{R_0}^R{r}^2.Math.\mathrm{dr}}$
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.

(19) 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 sixth section S6 situated at an eighth 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 0.87 to 1.27, which represents variation of 20% about the mean chord c. The chord at the blade start is equal to 0.37.

(20) Thereafter, the chords of the profiles 15 of the sections of the blade 1 are greater than the mean chord c between the sixth section S6 and a seventh section S7 situated at a ninth 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 seventh section S7 to the blade tip 9.

(21) In addition, the chord c decreases following a curve that is substantially parabolic beyond an eighth section S8 situated at a tenth distance equal to 0.95R. The end of the blade 1 thus forms a parabolic tip cap 8.

(22) The 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.

(23) 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.

(24) Initially, the twist angle varies little 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 0.35R. The variation in the twist angle is less than 2 between the start 3 of the airfoil portion 4 and the second section S2. 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 second section S2.

(25) Thereafter, the twist angle decreases between the second section S2 and a third section S3 situated at a third 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 third section S3.

(26) Thereafter, the twist angle decreases less between the third section S3 and a fourth section S4 situated at a fourth 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 fourth section S4. 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 0.7R.

(27) The twist angle again decreases more between the fourth section S4 and a fifth section S5 situated at a fifth 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 fifth section S5.

(28) Finally, the twist angle decreases between the fifth section S5 and the blade tip 9, the twist gradient increasing up to a twist gradient equal to 8/R at the blade tip 9.

(29) This twist relationship combined with the relationships for variation in the chord of the profiles 15 of the sections of the blade 1 serves to improve the aerodynamic performance of the blade 1, both during hovering flight and during forward flight.

(30) 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.

(31) Thus, the sweep is initially directed towards the front of the blade 1 between the start 3 of the airfoil portion 4 and a ninth section S9 situated at an eleventh 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 ninth section S9 and a tenth section S10 situated at a twelfth 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 tenth section S10 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.

(32) 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).

(33) Furthermore, the blade 1 has a downwardly-directed dihedral 5 at its free end. This dihedral 5 begins in the vicinity of the eighth section S8 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.

(34) 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.