JP2000118499A - Rotor blade for rotary wing plane - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a rotor blade for a rotary wing plane to reduce the generation of high speed impact noises. SOLUTION: A root part is mounted on a rotor head for rotational drive. A central part is linearly extended from the root part and its shape is specified by front and rear edges 21 and 22. The plan shape of a wing end part 12 is specified by a first front edge 23 protruded further frontward toward the outside from the outer end P1 of the front edge 21 of the central part 11, a second front edge 24 moved further backward toward the outside from the outer end P2 of the first front edge 23, and a side edge 25 and a rear edge 30. The backward movement angle of the rear edge 30 is higher than that of the second front edge 24, a chord length C1 between the second front edge 24 and the rear edge 30 is increased toward the outside. This constitution suppresses the generation of an impact wave and reduces the generation of high speed impact noises.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a rotor blade of a rotary wing machine such as a helicopter.

[0002]

FIG. 17 is a diagram showing rotor aerodynamic characteristics when a helicopter flies forward. As shown in FIG. 17A, when the helicopter 1 in which the rotor blade of the rotor radius R rotates at the angular velocity Ω moves forward at the flight speed V, the forward movement becomes a state where the flight speed V is added to the rotor speed ΩR. The airspeed is significantly different between the side blade and the retreating blade in which the flight speed V is subtracted from the rotor speed ΩR.

At the position of azimuth angle (counterclockwise angle with respect to the rear of the helicopter 1) Ψ = 90 °, the airspeed of the forward blade is maximized, and the airspeed at the tip of the blade is ΩR + V. . On the other hand, at the position where the azimuth angle Ψ = 270 degrees, the airspeed of the retreating blade becomes minimum, and the airspeed at the blade tip becomes ΩR-V. Furthermore, the airspeed at the intermediate position of the blade is a value obtained by proportionally distributing ΩR + V and ΩR−V. For example, ΩR = 795 km /
Assuming that h and V = 278 km / h, the airspeed becomes zero at a position about 35% from the root of the retreating blade, as shown in FIG.

[0004] Especially when the helicopter flies at high speed,
At the tip of the forward blade, the airspeed becomes transonic,
Strong shock waves are generated. This strong shock wave increases the resistance of the blade and generates strong shock noise called high-speed shock noise. At this time, in the coordinate system viewed from the rotating rotor blade, a phenomenon called non-localization of the supersonic region occurs. The generated shock wave propagates to a distant place through a non-localized supersonic region, and the noise audible to a distant place increases.

[0005] Since the airspeed of the retreating blade is greatly reduced, it is necessary to increase the angle of attack α of the blade in order to obtain the same lift as the advancing blade. Generally, the pitch angle of the blade is reduced to the azimuth angle. The pitch control is performed according to Ψ. The pitch angle of the blade is controlled by a sine wave which is minimum at azimuth angle Ψ = 90 degrees and maximum at ア ジ = 270 degrees, and the attack angle α of the blade at that time.
FIG. 17 shows the flapping motion of the blade itself.
As shown in (b), it changes in the span direction. For example, when Ψ = 90 degrees, the angle of attack α of the blade is about 0 degrees at the root and about 4 degrees at the tip. At Ψ = 270 degrees, the angle of attack α of the blade is about 0 degrees at the root and about 16 degrees at the tip.
It becomes 18 degrees, exceeding the stall angle. When the angle of attack α exceeds the stall angle, the lift coefficient Cl and the pitching moment coefficient Cm change suddenly, which leads to a large body vibration and the generation of a fatigue load on the pitch link.

As described above, the evaluation item of the forward blade is high-speed impact noise, and the evaluation items of the backward blade are the maximum lift coefficient Clmax and the stall angle. Note that the maximum lift coefficient Clmax is defined by the maximum value of the lift coefficient Cl at this time when the angle of attack α of the blade having a predetermined airfoil is gradually increased and the angle of attack α reaches the stall angle.
Generally, the smaller the absolute value of the high-speed impact noise and the pitching moment coefficient Cm, and the maximum lift coefficient Clma
The greater x and stall angle, the better the blade.

[0007]

As a method of reducing the high-speed impact noise, there is a method in which a swept angle is provided at the blade tip. By setting the sweepback angle, the shock wave is weakened and the noise is slightly reduced, but the noise is still large. If the swept angle is further increased, the tip stall occurs at a small angle of attack, and the maximum lift coefficient Clmax decreases.

Also, the stall angle and the maximum lift coefficient Clma
In order to improve x, there is a method of providing a protrusion at the leading edge of the blade tip of the rotor blade of the helicopter. The overhang is sometimes referred to as Dog Tooth. Near the overhang of the rotor blade, a vertical vortex is generated by the airflow from the front of the blade, and the flow on the blade surface is activated, but if the angle of attack is further increased,
Separation of the airflow occurs from the rear inside the overhang, resulting in a stall condition.

An object of the present invention is to provide a rotor blade of a rotary wing machine capable of suppressing the generation of shock waves and reducing high-speed impact noise.

Another object of the present invention is to provide a rotor blade of a rotary wing aircraft having good flight performance by increasing the stall angle and the maximum lift coefficient.

[0011]

SUMMARY OF THE INVENTION The present invention is directed to a root portion 9 attached to a rotor head for rotational driving, a central portion 11 extending linearly from the root portion 9, and a radially outward portion extending from the central portion 11. The first leading edge 23 extending, the second leading edge 24 extending outward while retreating from the outer end P2 of the first leading edge 23, and retreating from the outer end P3 of the second leading edge 24 to the wing tip P4. Side edge 25 extending outward, and wing tip P4 from central portion 11
A wing tip 12 having a shape defined by a trailing edge 30 extending outwardly to the outside, wherein the chord length C1 between the second leading edge 24 and at least a portion of the trailing edge 30 is longer toward the outside. This is a rotor blade of a rotary wing machine.

According to the present invention, by increasing the chord length C1 outside the wing tip 12, the blade thickness ratio can be reduced, so that the generation of shock waves can be suppressed and high-speed impact noise can be reduced. Can be.

The present invention also provides a root portion 9 attached to a rotor head for rotational driving, a central portion 11 extending linearly from the root portion 9, and a first front portion extending outwardly while advancing from the central portion 11. The edge 23, the second front edge 24 extending outward while retreating from the outer end P2 of the first front edge 23, the second front edge 24
A side edge 25 extending outward while retracting from the outer end P3 to the wing tip P4, and a wing end portion 12 having a shape defined by a trailing edge 30 extending outward from the central portion 11 to the wing tip P4, Near the outer end of the central portion 11 is a rotor blade of a rotary wing aircraft, wherein the angle of attack is locally small.

According to the present invention, the vicinity of the root at the inner side of the overhang is a portion where the air flow is likely to separate at the rear edge behind the overhang, and by providing a locally large torsion angle, it is possible to provide a more local area than the others. Since the angle of attack is reduced, separation of the airflow is less likely to occur in this portion, the stall angle can be increased, and the maximum lift coefficient can be increased.

The present invention also relates to an outer end P of the first leading edge 23.
The average aerodynamic center outside the center 2 is the leading edge 2 of the central portion 11.
It is characterized by being located at least 25% of the chord length C from 1 behind.

According to the present invention, the outer end P2 of the first leading edge 23
The outer average center of aerodynamics is generally further behind the feathering axis located 25% behind the chord length C from the leading edge 21 of the central portion 11. Therefore, when the angle of attack is large and the lift of the wing tip is large, the wing tip generates a pitching moment that reduces the angle of attack around the feathering axis, so that the blade twists in the direction of decreasing the angle of attack of the blade. Thereby, the blade tip stall can be suppressed.

Further, the present invention is characterized in that a droop is formed near the first front edge 23.

According to the present invention, by providing a droop near the first leading edge 23, the lift coefficient at the stall angle can be further increased, and the flight performance of a rotary wing aircraft such as a helicopter can be improved.

The present invention also relates to the wing tip 12 having a trailing edge 30.
A first trailing edge 26 adjacent to and behind the second leading edge 24
And a second trailing edge 27, wherein the first trailing edge 26 has a smaller receding angle than the second leading edge 24, and the second trailing edge 27 has a greater receding angle than the second leading edge 24. I do.

According to the present invention, the chord length C1 from the second leading edge 24 to the first trailing edge 26 becomes shorter toward the outside, and the second leading edge 2
The chord length C1 from 4 to the second trailing edge 27 is longer toward the outside. As described above, by increasing the chord length C1 of the outer portion of the second leading edge 24 toward the outside, high-speed impact noise can be reduced.

Since the chord length C1 from the second leading edge 24 to the first trailing edge 26 is shorter toward the outside, the area of the wing tip 12 can be made relatively small. Therefore, the resistance of the rotor blade can be reduced, the lift-drag ratio can be improved, and the power required for hovering can be reduced, that is, hovering performance can be improved.

The present invention also relates to the wing tip 12 having a trailing edge 30.
Further includes a third trailing edge 28 extending from the outer end P7 of the second trailing edge 27 to the wing tip P4 and having a smaller receding angle than the second trailing edge 27.

According to the present invention, as described above, the chord length C1 of the inner portion of the second leading edge 24 is made shorter toward the outer side, thereby suppressing an increase in the area of the wing tip and reducing the resistance. The high-speed impact noise can be reduced by increasing the chord length C1 of the outer portion of the second leading edge 24 toward the outside. Further, by providing a third trailing edge 28 having a smaller receding angle than the second trailing edge 27 outside the second trailing edge 27, the aerodynamic center is adjusted so as not to be excessively rearward, and
Sufficient strength can be ensured near the wing tip P4.

Further, according to the present invention, an outer end P of the first leading edge 23 is provided.
2 is 85% to 9 of the blade length R from the rotation center of the rotor.
It is characterized by being located 0% outside.

According to the present invention, the outer end P of the first leading edge 23
2. In other words, the peak of the overhang is arranged 85% to 90% outside of the blade length R from the rotation center of the rotor where the shock wave is likely to be generated, so that the generation of the shock wave is reliably suppressed and the high-speed shock noise is reduced. Can be.

Further, according to the present invention, an outer end P of the second front edge 24 is provided.
3 is 93% to 9 of the blade length R from the rotation center of the rotor.
It is characterized by being located 6% outside.

According to the present invention, the second leading edge 24 having a relatively small swept angle is set at a blade length R of 9 from the rotation center of the rotor.
By setting it to 3% to 96% outside, the side edge 25 having a relatively large swept angle can be provided, and the occurrence of blade tip stall can be reliably suppressed.

[0028]

(First Embodiment) FIG. 1 is a plan view showing a first embodiment of the present invention. Rotor blade 1
0 comprises a root 9, a center 11 and a wing tip 12. The root portion 9 is a member attached to a rotor head that drives the rotor blade 10 to rotate. The central part 11
It extends linearly from the root 9 and occupies most of the rotor blade 10. Most of the aerodynamic characteristics of the rotor blade 10 are determined by the airfoil shape, which is the cross-sectional shape of the central portion 11, and the airfoil shape viewed from above. The shape of the central portion 11 when viewed from above is a front edge 21 and a rear edge 2 which are parallel to each other.
2, the chord length C of the central portion 11 is equal to the leading edge 2
1 and the distance of the trailing edge 22. Wing tip 1
Reference numeral 2 denotes a tip portion of the rotor blade 10 extending from the central portion 11 to the opposite side of the root portion 9.

The rotor blade 10 is a main wing of the helicopter, is driven by the rotor head, and makes the helicopter fly by rotating the wing tip 12 outward so as to draw a disk.

FIG. 2 is a partially enlarged view showing the shape of the rotor blade 10 of FIG. The shape of the wing tip 12 is defined by a first leading edge 23, a second leading edge 24, a side edge 25 and a trailing edge 30. The first leading edge 23 is a leading edge 21 of the central portion 11.
Of the first front edge 23, and extends to the outer end P2 of the first front edge 23. Here, the term “outside” indicates a blade tip side in the direction Y which is the span direction of the rotor blade 10, and the outer end indicates an end point on the blade tip side in the direction Y.

The outer end P2 is separated from the center of rotation of the rotor blade 10 by a distance of 0.85R to 0.90R,
The front edge 21 is disposed forward from the front edge 21 by a distance of 0.3C to 0.45C. Here, R is the blade length, that is, the length of the rotor blade 10 in the span direction, and C is FIG.
Similarly, the distance between the front edge 21 and the rear edge 22 of the central portion 11. In FIG. 2, the length of each part is entered, and a representative value is entered in parentheses. For example, a representative value of the distance between the outer end P2 and the rotation center of the rotor blade 10 is 0.8
8R.

The second leading edge 24 is provided at the outer end P2 of the first leading edge 23.
From the outside to the outside, the second leading edge 2
4 to the outer end P3. The sweepback angle of the second leading edge 24 is 15 to 20 degrees. The outer end P3 is arranged at a distance of 0.93R to 0.96R from the center of rotation of the rotor blade 10.

As the side edge 25 goes outward from the outer end P3 of the second leading edge 24, it retreats backward so as to draw a parabola and extends to the wing tip P4. Also, the side edge 25
Side edge 25a extending straight to wing tip P4 at a sweepback angle of 70 degrees
It may be.

The trailing edge 30 retreats rearward from the outer edge P5 of the trailing edge 22 of the central portion 11 to the outside, and the wing tip P
It extends to four. The outer end P5 has a distance from the center of rotation of the rotor blade 10 of 0.85R to 0.
They are located 90R apart. The receding angle of the trailing edge 30 is 2
5 degrees to 30 degrees, and the chord length C1 between the second leading edge 24 and the trailing edge 30 is longer toward the outside.

Next, the aerodynamic center of the rotor blade 10 will be described. First, in FIGS. 1 and 2, the origin of the XY coordinates is taken as the rotation center of the rotor blade 10, and the Y axis is made to coincide with the feathering axis. The X axis has a forward direction behind the rotor blade 10, and the Y axis has a rotor blade 1
The wing tip side of 0 is the positive direction. The feathering axis is an axis around which the rotor blade 10 rotates when the angle of attack is adjusted, and is usually 0.25 from the front edge 21 of the central portion 11.
C is located behind.

The outer end P2 of the first leading edge 23 is set to 0.
Assuming that the aerodynamic center is located at the position of 88R, the average aerodynamic center from the distance of zero to the distance of 0.88R is located 0.25C behind the front edge 21 of the central part 11, similarly to the feathering axis. The average aerodynamic center outside the outer end P1 of the first front edge 23 is 0.2 mm from the front edge 21 of the central portion 11.
It is located further back than the position behind 5C. The trajectory drawn by the second leading edge 24 and the side edge 25 in the XY coordinate system is represented by X =
F (Y), the trajectory of the trailing edge 30 is X = G (Y), and the aerodynamic center position in each section is X = H (Y) (≒ 0.75 · F (Y) +
0.25 · G (Y)), the outer end P of the first front edge 23
The average aerodynamic center outside of 1 is located a distance Z behind the feathering axis. Where Z is the integration interval 0.8
{H (Y)} G (Y) -F (Y)} d at 8R to 8R
Y / {G (Y) -F (Y)} dY. Z is
For example, it is 0.18C.

As described above, since the aerodynamic center of the wing tip portion outside of P2 is located further rearward than the feathering axis, when the attack angle of the blade is large and the lift of the wing tip is large, the feathering axis Since a pitching moment that reduces the angle of attack is generated around the blade, the blade is twisted in a direction to reduce the angle of attack, and the blade tip stall can be suppressed.

FIG. 3 is a graph showing the torsional angle with respect to the distance from the rotation center in the first embodiment. The horizontal axis of the graph indicates the distance from the rotation center of the rotor, and the vertical axis indicates the torsional angle. As shown in the graph, in the rotor blade 10 of the first embodiment, the pitch angle decreases linearly toward the outer side. At the position of the innermost distance of zero, the pitch angle is 8 degrees, and at the position of the outermost distance R, the pitch angle is -2.5 degrees.

FIG. 4 is a plan view showing the shape of a rotor blade 51 as a first comparative example. The shape of the rotor blade 51 of the first comparative example is the same as that of the rotor blade 1 of the first embodiment.
0, except that the trailing angle of the trailing edge 30 is
The second leading edge 24 and the trailing edge 3 are smaller than the retreat angle of the leading edge 24.
The chord length CI between 0 and P is substantially constant between P2 and P3.

FIG. 5A shows the distribution of the Mach number (ratio of flow velocity to sound velocity) around the rotor blade of the second comparative example.
(B) shows the Mach number distribution around the rotor blade of the first comparative example, and FIG. 5 (c) shows the Mach number distribution around the rotor blade of the first embodiment.

In the second comparative example shown in FIG. 5A, a phenomenon called non-localization of a supersonic region in which the supersonic region on the rotor blade surface is connected to a distant supersonic region has occurred. In the first embodiment c), the intensity of the shock wave is weakened by the shape effect of the first embodiment, and non-localization of the supersonic region does not occur. Also, from the comparison between the first comparative example of FIG. 5B and the first embodiment of FIG. 5C, the reverse taper of the first embodiment (chord length between the second leading edge 24 and the trailing edge 30). C
By increasing the length of the blade toward the outside, the blade thickness ratio at the blade tip is reduced, the strength of the shock wave on the blade surface is reduced, and the pressure fluctuation near the blade is also reduced.

FIG. 6A shows the sound pressure fluctuation of the second comparative example.
6B is a graph showing the sound pressure fluctuation of the first comparative example, and FIG. 6C is a graph showing the sound pressure fluctuation of the first embodiment. FIG.
The minimum value of the sound pressure of the second comparative example shown in FIG.
a, the minimum value of the sound pressure of the first comparative example shown in FIG.
While the pressure is 40 Pa, the minimum value of the sound pressure of the first embodiment shown in FIG. 6C is -190 Pa, which indicates that the noise of the first embodiment is reduced.

In conclusion, the reverse tapered shape of the first embodiment shown in FIG. 5 has a high-speed impact noise shown in the noise waveform of FIG. 6 due to suppression of non-localization in the supersonic region and weakening of the shock wave near the blade. To reduce.

(Second Embodiment) FIG. 7 is a plan view showing the shape of a rotor blade 62 according to a second embodiment. Since the shape of the rotor blade 62 is the same as that of the first embodiment, the description is omitted. However, a local torsion is formed in a region A near the outer end of the central portion 11.

FIG. 8 is a graph showing the torsional angle with respect to the distance from the rotation center. As in FIG. 3, the rotor blades 62 are twisted so that the pitch angle becomes smaller as going outward. The change in pitch angle is generally linear, that is, linear. Furthermore, near the outer edge of the center, a local torsion is formed,
That is, the pitch angle fluctuates in a curve. In the area A, the pitch angle suddenly decreases and then increases again to return to the original pitch angle.

FIG. 9 shows a rotor blade 52 of a second comparative example.
It is a top view which shows the shape of. The blade tip of the rotor blade 52 of the second comparative example is rectangular.

FIG. 10 shows a second comparative example and first, second, and third comparative examples.
It is a graph which shows the angle-of-attack dependence of lift count Cl of a 3rd embodiment. The horizontal axis of the graph is the angle of attack of the blade (unit is de
g), and the vertical axis indicates the lift coefficient Cl. The rotor blade of the first embodiment has a larger stall angle and a maximum lift count Clmax than the rotor blade of the second comparative example, but a drop in the lift temporarily occurs near the attack angle of 15 °. As shown by the vector distribution in FIG. 11 (a) and the surface streamline in FIG. 11 (b), this drop in lift is caused by the separation of the flow from the trailing edge of the leading edge of the blade. . Note that the lift characteristics of the first comparative example are almost the same as the lift characteristics of the first embodiment, and a similar problem of flow separation at the trailing edge occurs. In the second embodiment of FIG. 10, this separation is suppressed by local torsion, and the sudden drop in the lift coefficient of the first embodiment is improved.
The rotor blade of the third embodiment will be described later.

(Third Embodiment) FIG. 13 shows a third embodiment of the present invention.
It is a partial plan view showing the shape of the rotor blade 63 of the embodiment. First Leading Edge 23 of Rotor Blade of Second Embodiment
A droop is provided in the vicinity. The droop is provided near the leading edge, and is a portion of the airfoil that sags as it goes forward. An example of droop is shown in FIG.

As shown in FIG. 10, the third embodiment is the second embodiment.
The lift coefficient is further improved than in the embodiment.

(Fourth Embodiment) FIG. 14 shows a fourth embodiment of the present invention.
FIG. 3 is a partial plan view illustrating a shape of a rotor blade 64 of the embodiment. The shape of the wing tip 12 is defined by a first leading edge 23, a second leading edge 24, a side edge 25, a first trailing edge 26 and a second trailing edge 27. The first leading edge 23 is the leading edge 2 of the central portion 11.
1 as it goes outward from the outer end P1,
The first front edge 23 extends to the outer end P2. The outer end P2 is
Distance 0.85R from the center of rotation of rotor blade 64
It is located 0.90R away from the front edge 21 of the central portion 11 and is located forward by a distance of 0.3C to 0.45C.

The second leading edge 24 is located at the outer end P2 of the first leading edge 23.
From the outside to the outside, the second leading edge 2
4 to the outer end P3. The sweepback angle of the second leading edge 24 is 15 to 20 degrees. The outer end P3 is arranged at a distance of 0.93R to 0.96R from the rotation center of the rotor blade 64.

As the side edge 25 goes outward from the outer end P3 of the second leading edge 24, it retreats backward so as to draw a parabola and extends to the wing tip P4. Also, the side edge 25
The shape may extend linearly to the blade tip P4 at a sweepback angle of 70 degrees.

The first rear edge 26 linearly extends outward from the outer edge P5 of the rear edge 22 of the central portion 11 to the outer edge P6.
The outer end P5 is disposed at a distance of 0.85R to 0.90R from the center of rotation of the rotor blade 64, like the outer end P2. The second trailing edge 27 retreats backward from the outer end P6 of the first trailing edge 26 to the outside, and extends to the wing tip P4. The sweepback angle of the second trailing edge 27 is 30 degrees to 45 degrees.
Degrees. Therefore, the chord length C1 between the second leading edge 24 and the first trailing edge 26 becomes shorter toward the outside until P6,
On the outside of P6, the chord length C1 between the second leading edge 24 and the second trailing edge 27 is longer toward the outside.

With respect to the center of aerodynamic force of the rotor blade 64, as in the first embodiment, the distance from zero to 0.88
The average aerodynamic center up to R is the same as the feathering axis,
It is located 0.25C behind the front edge 21 of the central portion 11. The average aerodynamic center outside the outer end P1 of the first front edge 23 is located further rearward than the position 0.25C behind the front edge 21 of the central portion 11.

Further, the rotor blade 6 of the fourth embodiment
4, as in the first embodiment shown in FIG.
The pitch angle decreases linearly toward the outside, and a local torsion is formed.

FIG. 15 is a graph showing the lift-drag ratios of the first comparative example and the first and fourth embodiments. The horizontal axis of the graph shows the resistance coefficient, and the vertical axis shows the lift coefficient. The lift-drag ratio is the ratio of the lift coefficient to the drag coefficient. The rotor blade 64 of the fourth embodiment is different from the rotor blade 61 of the first embodiment.
The lift-drag ratio is larger than that. This is the second leading edge 2
4 shows that the chord length C1 between the first trailing edge 26 and the first trailing edge 26 is made shorter toward the outside, and the area of the wing tip 12 is reduced, thereby increasing the lift-drag ratio.

As shown in FIG. 15, in a normal helicopter, hovering is performed under conditions such that the average lift coefficient is in the range of 0.4 to 0.5. Since the rotor blade 64 of the fourth embodiment also performs hovering in the above-described lift coefficient range, hovering can be performed with less power than in the first embodiment and the first comparative example.

(Fifth Embodiment) FIG. 16 shows a fifth embodiment of the present invention.
It is a partial plan view showing the shape of the rotor blade 65 of the embodiment. The rotor blade 65 has a shape obtained by removing the tip 31 from the rotor blade 64 of the fourth embodiment. The shape of the wing tip 12 includes a first leading edge 23, a second leading edge 24, a side edge 25, a first trailing edge 26, a second trailing edge 27, and a third trailing edge 28.
Defined by The first front edge 23 extends forward from the outer edge P1 of the front edge 21 of the central portion 11 toward the outside, and extends to the outer edge P2 of the first front edge 23. Outer end P2
Is 0.85R from the center of rotation of the rotor blade 65.
It is arranged apart from the front edge 21 of the central portion 11 by a distance of 0.3C to 0.45C and separated by .about.0.90R.

The second leading edge 24 is the outer end P2 of the first leading edge 23.
From the outside to the outside, the second leading edge 2
4 to the outer end P3. The sweepback angle of the second leading edge 24 is 15 to 20 degrees. The outer end P3 is arranged at a distance of 0.93R to 0.96R from the rotation center of the rotor blade 65.

As the side edge 25 goes outward from the outer end P3 of the second leading edge 24, it retreats backward so as to draw a parabola and extends to the wing tip P4. Also, the side edge 25
The shape may extend linearly to the blade tip P4 at a sweepback angle of 70 degrees.

The first rear edge 26 linearly extends outward from the outer edge P5 of the rear edge 22 of the central portion 11 to the outer edge P6.
The outer end P5 is spaced apart from the center of rotation of the rotor blade 65 by a distance of 0.85R to 0.90R, like the outer end P2, and the second trailing edge 27 is located outward from the outer end P6 of the first trailing edge 26. As it moves, it retreats rearward and extends to the outer end P7. The sweepback angle of the second trailing edge 27 is 30 degrees to 45 degrees. Therefore, the chord length C1 between the second leading edge 24 and the first trailing edge 26 becomes shorter toward the outside up to P6, and P
6, the chord length C1 between the second leading edge 24 and the second trailing edge 27 is longer toward the outside.

The third trailing edge 28 is the outer end P7 of the second trailing edge 27.
As it goes outward from the wing, it retreats backward, and the wing tip P4
Extending to The sweepback angle of the third trailing edge 28 is between 0 degree and 20 degrees.
Degrees. By providing the third trailing edge 28 having a sweepback angle smaller than the second trailing edge 27, the aerodynamic center is adjusted so as not to be excessively rearward, and the vicinity of the wing tip P4 is the fourth embodiment. It is slightly thicker than, ensuring strength.

As in the first embodiment, the center of aerodynamic force of the rotor blade 65 is changed from zero distance to 0.88 distance.
The average aerodynamic center up to R is the same as the feathering axis,
It is located 0.25C behind the front edge 21 of the central portion 11. The average aerodynamic center outside the outer end P1 of the first front edge 23 is located further rearward than the position 0.25C behind the front edge 21 of the central portion 11, but the third rear edge 28
Can be adjusted by appropriately selecting the angle.

Further, the rotor blade 65 of the fifth embodiment
Also, as in the first embodiment shown in FIG. 3, the pitch angle decreases linearly toward the outside, and a local torsion is formed.

[0065]

As described above, according to the present invention, the wing tip 1
By increasing the chord length C1 on the outer side, the blade thickness ratio can be reduced, so that the generation of shock waves can be suppressed and high-speed impact noise can be reduced.

Further, according to the present invention, by lowering the twist near the root at the inner side of the overhang locally than in the others, vortex separation hardly occurs in this portion, and the stall angle can be increased. , The maximum lift coefficient can be increased.

According to the present invention, the average aerodynamic center outside the outer end P2 of the first front edge 23 is generally adjusted to the center 11
By arranging further behind the feathering axis located 25% of the chord length C from the leading edge 21, a pitching moment that reduces the angle of attack about the feathering axis is generated, and the tip stall occurs. Can be suppressed.

Further, according to the present invention, by providing the droop, the lift coefficient can be further increased and the flight performance can be improved.

Further, according to the present invention, by shortening the chord length C1 of the inner portion of the second leading edge 24, the area of the wing tip can be reduced and the hovering characteristics can be improved. By increasing the chord length C1 of the outer portion of the leading edge 24, high-speed impact noise can be reduced.

According to the present invention, the second trailing edge 24
The third trailing edge 2 having a smaller receding angle than the second trailing edge 27
By providing 8, the aerodynamic center can be adjusted to be located too rearward, the angle between the side edge 25 and the trailing edge 30 can be increased, and sufficient strength can be secured near the wing tip P4.

According to the present invention, the top of the overhang is 85% of the blade length from the center of rotation of the rotor blade 65.
By disposing 90% outside, it is possible to reliably suppress the generation of shock waves and reduce high-speed impact noise.

Further, according to the present invention, the second leading edge 24 is set at 93% to 9% of the blade length from the rotation center of the rotor blade 65.
By setting it to 6% outside, the side edge 25 having a relatively large swept angle can be provided, and the occurrence of blade tip stall can be reliably suppressed.

[Brief description of the drawings]

FIG. 1 is a plan view showing a first embodiment of the present invention.

FIG. 2 is an enlarged plan view showing the shape of the rotor blade 10 of FIG.

FIG. 3 is a graph showing a torsion angle with respect to a distance from a rotation center in the first embodiment.

FIG. 4 is a plan view showing the shape of a rotor blade 51 of a first comparative example.

5A is a diagram illustrating a second comparative example, FIG. 5B is a diagram illustrating a first comparative example, and FIG. 5C is a diagram illustrating a Mach number distribution near a blade tip of the first embodiment. .

FIG. 6A is a graph showing a sound pressure fluctuation caused by a rotor blade according to a second comparative example, and FIG.
FIG. 6C is a graph illustrating a high-pressure fluctuation caused by the rotor blade 51 of the comparative example, and FIG. 6C is a graph illustrating a sound pressure fluctuation caused by the rotor blade 10 of the first embodiment.

FIG. 7 is a plan view illustrating a shape of a rotor blade 62 according to a second embodiment of the present invention.

FIG. 8 is a graph showing a torsion angle with respect to a distance from a rotation center in the second embodiment.

FIG. 9 is a plan view illustrating a shape of a rotor blade 52 of a second comparative example.

FIG. 10 is a graph showing the angle-of-attack dependence of the lift coefficient Cl of the second comparative example, and the angle-of-attack dependence of the lift coefficient Cl of the first, second, and third embodiments.

FIG. 11 is a view showing separation of a flow generated at a rear edge portion of the protruding inner portion which occurs in the first embodiment.

FIG. 12 is a diagram illustrating an example of droop.

FIG. 13 shows a rotor blade 63 according to a third embodiment of the present invention.
It is a top view which shows the shape of.

FIG. 14 shows a rotor blade 64 according to a fourth embodiment of the present invention.
It is a top view which shows the shape of.

FIG. 15 is a graph showing the lift-drag ratios of the first comparative example and the second and fourth embodiments.

FIG. 16 shows a rotor blade 65 according to a fifth embodiment of the present invention.
It is a top view which shows the shape of.

FIG. 17 is a diagram illustrating rotor aerodynamic characteristics when the helicopter flies forward.

[Explanation of symbols]

 REFERENCE SIGNS LIST 9 root portion 10 rotor blade 11 central portion 12 wing tip 21 leading edge of central portion 22 trailing edge of central portion 23 first leading edge of wing tip 24 second leading edge of wing tip 25 side edge of wing tip 26 First trailing edge of wing tip 27 Second trailing edge of wing tip 28 Third trailing edge of wing tip 30 Trailing edge of wing tip P1 Outer edge of leading edge at center P2 First wing tip Outer edge of leading edge P3 Outer edge of second leading edge of wing tip P4 Wing tip P5 Outer edge of trailing edge in center P6 Outer edge of first trailing edge of wing tip P7 Second trailing edge of wing tip Outer end C, C1 chord length

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[Procedure amendment]

[Submission date] September 27, 1999 (September 9, 1999
7)

[Procedure amendment 1]

[Document name to be amended] Statement

[Correction target item name] Claims

[Correction method] Change

[Correction contents]

[Claims]

[Procedure amendment 2]

[Document name to be amended] Statement

[Correction target item name] 0009

[Correction method] Deleted

[Procedure amendment 3]

[Document name to be amended] Statement

[Correction target item name] 0010

[Correction method] Change

[Correction contents]

It is an object of the present invention to provide a rotor blade of a rotary wing aircraft having good flight performance by increasing the stall angle and the maximum lift coefficient.

[Procedure amendment 4]

[Document name to be amended] Statement

[Correction target item name] 0011

[Correction method] Change

[Correction contents]

[0011]

SUMMARY OF THE INVENTION The present invention is directed to a root portion 9 attached to a rotor head for rotational driving, a central portion 11 extending linearly from the root portion 9, and a radially outward portion extending from the central portion 11. The first leading edge 23 extending, the second leading edge 24 extending outward while retreating from the outer end P2 of the first leading edge 23, and retreating from the outer end P3 of the second leading edge 24 to the wing tip P4. Side edge 25 extending outward, and wing tip P4 from central portion 11
A wing tip 12 having a shape defined by a trailing edge 30 extending outwardly to the outside, and a local torsion reduction is formed near an outer end of the central portion 11. It is a rotor blade.

[Procedure amendment 5]

[Document name to be amended] Statement

[Correction target item name] 0012

[Correction method] Change

[Correction contents]

According to the present invention, the vicinity of the root at the inner side of the overhang is a portion where the air flow is likely to separate at the rear edge behind the overhang, and by providing a locally large torsion angle, it is possible to provide a more local area than the others. Since the angle of attack is reduced, separation of the airflow is less likely to occur in this portion, the stall angle can be increased, and the maximum lift coefficient can be increased.

[Procedure amendment 6]

[Document name to be amended] Statement

[Correction target item name] 0013

[Correction method] Change

[Correction contents]

The present invention also relates to the second leading edge 24 and the trailing edge 3.
The chord length C1 between at least a part of 0 and the outside is characterized by being longer toward the outside.

[Procedure amendment 7]

[Document name to be amended] Statement

[Correction target item name] 0014

[Correction method] Change

[Correction contents]

According to the present invention, by increasing the chord length C1 outside the wing tip 12, the blade thickness ratio can be reduced, so that the generation of shock waves can be suppressed and high-speed impact noise can be reduced. Can be.

[Procedure amendment 8]

[Document name to be amended] Statement

[Correction target item name] 0026

[Correction method] Change

[Correction contents]

Further, according to the present invention, an outer end P of the second front edge 24 is provided.
3 is 93% to 9 of the blade length R from the rotation center of the rotor.
6% outside, the side edge 25 has a receding angle of 60 to 75 degrees.

[Procedure amendment 9]

[Document name to be amended] Statement

[Correction target item name] 0027

[Correction method] Change

[Correction contents]

According to the present invention, the second leading edge 24 having a relatively small swept angle is set at a blade length R of 9 from the rotation center of the rotor.
By setting the receding angle of the side edge 25 to 60 degrees to 75 degrees, the side edge 25 having a relatively large receding angle can be provided, and the occurrence of blade tip stall can be ensured. Can be suppressed.

[Procedure amendment 10]

[Document name to be amended] Statement

[Correction target item name] 0065

[Correction method] Change

[Correction contents]

[0065]

As described above, according to the present invention, the twisting near the inner root of the overhang is locally increased more than the others, so that vortex separation hardly occurs in this portion, and the stall angle is reduced. The maximum lift coefficient can be increased.

[Procedure amendment 11]

[Document name to be amended] Statement

[Correction target item name] 0066

[Correction method] Change

[Correction contents]

Further, according to the present invention, the blade thickness ratio can be reduced by increasing the chord length C1 outside the blade tip 12, so that the generation of shock waves can be suppressed, and high-speed impact noise can be reduced. be able to.

[Procedure amendment 12]

[Document name to be amended] Statement

[Correction target item name] 0072

[Correction method] Change

[Correction contents]

Further, according to the present invention, the second leading edge 24 is set at 93% to 9% of the blade length from the rotation center of the rotor blade 65.
By setting the retraction angle of the side edge 25 to 60 degrees to 75 degrees, the side edge 25 having a relatively large retraction angle can be provided, and the occurrence of blade tip stall can be reliably suppressed. Can be.

 ──────────────────────────────────────────────────続 き Continuing from the front page (72) Inventor Atsushi Murashige 2 Kawasaki-cho, Kakamigahara-shi, Gifu Pref.

Claims (8)

[Claims] 1. A root portion 9 attached to a rotor head for rotational driving, a central portion 11 linearly extending from the root portion 9, and a first front edge 2 extending outward while advancing from the central portion 11.
3, a second leading edge 24 extending outward while retracting from the outer end P2 of the first leading edge 23, and a lateral edge extending outward while retracting from the outer end P3 of the second leading edge 24 to the wing tip P4. 25, and a trailing edge 30 extending outward from the central portion 11 to the wing tip P4.
And a wing tip portion 12 having a shape defined by the following formula: wherein a chord length C1 between the second leading edge 24 and at least a part of the trailing edge 30 is longer toward the outside. blade. 2. A root portion 9 attached to a rotor head for rotational driving, a central portion 11 extending linearly from the root portion 9, and a first front edge 2 extending outward from the central portion 11 while moving forward.
3, a second leading edge 24 extending outward while retracting from the outer end P2 of the first leading edge 23, and a lateral edge extending outward while retracting from the outer end P3 of the second leading edge 24 to the wing tip P4. 25, and a trailing edge 30 extending outward from the central portion 11 to the wing tip P4.
A rotor blade of a rotary wing machine, comprising: a wing tip portion 12 having a shape defined by the following formula; 3. An average aerodynamic center outside the outer end P2 of the first front edge 23 is located at least 25% of the chord length C behind the front edge 21 of the central portion 11. A rotor blade for a rotary wing machine according to claim 1 or 2. 4. The rotor blade of a rotary wing machine according to claim 1, wherein a droop is formed near the first leading edge. 5. The trailing edge 30 of the wing tip 12 includes a first trailing edge 26 and a second trailing edge 27 adjacent to each other behind a second leading edge 24, the first trailing edge 26 being a second trailing edge. The receding angle is smaller than the leading edge 24,
The rotor blade of a rotary wing machine according to claim 1 or 2, wherein the second trailing edge (27) has a larger sweepback angle than the second leading edge (24). 6. The trailing edge 30 of the wing tip 12 further extends from the outer end P7 of the second trailing edge 27 to the wing tip P4 to form a third trailing edge 28 having a smaller receding angle than the second trailing edge 27. The rotor blade of a rotary wing machine according to claim 4, wherein the rotor blade comprises: 7. The rotary wing machine according to claim 1, wherein an outer end P2 of the first leading edge 23 is located 85% to 90% outside a blade length R from a rotation center of the rotor. Rotor blades. 8. The rotary wing machine according to claim 1, wherein the outer end P3 of the second leading edge 24 is located 93% to 96% outside the blade length R from the rotation center of the rotor. Rotor blades.