JP2001239997A - Airfoil section for blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To intend for accelerating a helicopter and increasing a pay load, by providing an airfoil section for a blade having a low pitching moment coefficient, capable of having a high maximum lift, a high resistance divergence Mach number, and a high drag ratio and capable of reducing a structural load applied to a pitch angle control device. SOLUTION: This airfoil section for a blade is formed into an airfoil section computed by the numerical formula 1 by using a coefficient within ±3% of the coefficient indicated in the table 1. That is, a large front radius, a large camber provided on a part excluding a rear rim part, a flat shape on the upper surface and the lower surface of the center part of the airfoil section are provided in this airfoil section.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a blade wing applied to a rotor blade provided above a helicopter for generating a lift L for lifting the helicopter in the air and for generating a propulsion necessary for forward flight. About the type. The blade airfoil of the present invention can also be applied as a blade airfoil for a propeller of an aircraft.

[0002]

2. Description of the Related Art As a conventional airfoil for a blade used for a rotor blade of a helicopter or the like (hereinafter simply referred to as an airfoil), there is a National Advisory Committee.
tee for Aeronautics (NACA), and the NAC shown in FIG.
A00XX series, or airfoil N with airfoil coordinates shown in Table 3.
An NACA 230XX type airfoil such as ACA23012 is used. The last two digits XX of the codes indicating the system of the airfoil 01 are the chord length c of the airfoil 01.
% Of blade thickness ratio t, which is the ratio of the maximum blade thickness of airfoil 01 to
Is shown.

[0003]

(Equation 3)

[0004]

[Table 3]

Here, y is a blade thickness direction coordinate, x is a chord direction coordinate, c is a chord length, and t is a ratio of a maximum blade thickness of the airfoil to a chord length c, a blade thickness ratio (%). Are respectively shown. Also, the blade thickness ratio t
Is simply a coefficient and has no particular significance.

[0006] Of these airfoils 01, NACA00X
As shown in FIG. 5, the shape of the X-type airfoil on the upper surface 02 and the lower surface 03 of the airfoil 01 calculated by Expression 3 is symmetrically arranged with respect to the chord line CL. It becomes wing type 01.
Further, among these airfoils 01, the airfoil 01 of the NACA230XX system is, as apparent from the airfoil coordinates of NACA23012 shown in Table 3, the upper surface 02 of the airfoil from the chord line CL.
It is a cambered airfoil with a small camber projecting slightly to the side.

For this reason, as shown in FIG. 6, the magnitude of the rotational force around the aerodynamic center CP of the airfoil 01, which is a point 1/4 of the chord length from the leading edge 04 of the airfoil 01, is reduced. As shown in the conventional airfoil data shown in Table 4, the pitching moment coefficient C _m shown in FIG.
Both of the ACA230XX series airfoils are very small, near zero.

[0008]

[Table 4]

Therefore, in the blade 06 of the helicopter 07 having such an airfoil 01, the angle of attack α of the blade 06 described later, which is unique to the helicopter 07 in forward flight.
There is an advantage that the structural load applied to the pitch angle control device that frequently controls (pitch angle) can be reduced.

[0010] For this purpose, the airfoil proposed by each helicopter maker, for example, the VR type which is a more advanced airfoil or the OA type which is a second generation type airfoil shown in FIG. 4 described later. or aerofoil SC system, etc., similar to those developed on the basis of the airfoil 01 developed by these NACA, is a small airfoil of pitching moment coefficient C _m, is at present that is used .

That is, most of the airfoils 01 used for the blade 06 are generally made to have a relatively symmetrical airfoil shape in order to reduce the pitching moment M.

On the other hand, recently, helicopter 07
Increased maximum take-off weight, so-called increased payload,
Increasing flight speed has been promoted, and improving performance has become an urgent issue.

Next, the operation and control of the blade 06 used as the rotor blade of the helicopter 07 will be described. FIG. 7 is a plan view showing the operating environment of the blade 06 during the forward flight of the helicopter 07.

As shown in the figure, the blade 06 is normally rotated counterclockwise in plan view. Therefore, the blade 06 rotating rightward in the traveling direction of the helicopter 07 has a rotation speed R of the blade 06. Since the forward speed F of the helicopter 07 at the time of flight is added to the forward speed component, the airfoil 01 is exposed to a large airspeed V.

Conversely, the blade 0 rotating leftward in the traveling direction
In No. 6, since the speed at which the forward speed F is reduced from the reverse speed component of the blade 06 rotation speed R becomes the airspeed V, the airfoil 01 moves at a low airspeed V.

Therefore, in such an airfoil 01, if the angles of attack α of the left and right blades 06 are, for example, the same, the angle of attack due to the difference in lift L generated according to the magnitude of the airspeed V is increased. A balance occurs, and it becomes difficult to maintain the horizontal attitude of the helicopter 07.

In order to cancel the unbalance of the lift L and maintain the horizontal posture of the helicopter 07, the lift L of the blade 06 rotating on the right side in the traveling direction of the helicopter 07 is reduced. As shown in b), the blade 06 rotating on the right side has a small angle of attack α, and furthermore, the blade 06 rotating at a position orthogonal to the flight direction.
In this case, since the airspeed V becomes the largest, it is necessary to change the angle of attack α in accordance with the rotation angle so that the angle of attack α becomes the smallest.

Further, in order to increase the lift L of the blade 06 rotating on the left side, as shown in FIG. 7C, the angle of attack α of the blade 06 is increased, and the blade 06 rotates at a position orthogonal to the flight direction. In the blade 06, since the airspeed V becomes the smallest, it is necessary to change the angle of attack α according to the rotation angle so that the angle of attack α becomes the largest. Like this
Pitch angle control that cyclically changes the attack angle α of the blade 06 according to the rotation position of the blade 06 during one rotation of the blade 06 is essential.

Therefore, as described above, the airfoil 01 of the blade 06 used for the rotor blade of the helicopter 07 has a pitching moment M in order to reduce the structural load applied to the pitch angle control device of the blade 06.
It is important to reduce.

In order to increase the payload or improve the forward flight performance required of the helicopter 07 in recent years,
The airfoil 01 is required to achieve the following performance.

(1) Realization of a high-resistance divergence Mach number for suppressing a rapid increase in resistance D during high-speed flight.

That is, when the airspeed V of the flying object (airfoil 01) moving in the atmosphere approaches the sound speed, a shock wave is generated on the surface of the airfoil 01, and the resistance D rapidly increases. The Mach number M, which is the ratio between the airspeed V and the sound speed at which the resistance D sharply increases, is referred to as a resistance divergence Mach number _{Mdd. The} Mach number M is plotted on the horizontal axis, and the resistance D is dimensionlessly plotted on the vertical axis. When the resistance coefficient C _d is taken, the generation of the resistance divergence Mach number M _dd defined by the Mach number M when the slope dC _d / dM of the curve becomes 0.1 is generated at the largest possible Mach number M. The realization of a high-resistance divergence Mach number, which is to be performed, is required to improve the forward flight performance when the helicopter 07 is flown at high speed.

(2) The lift for generating a large lift despite the decrease in the airspeed V, such as the blade 06 rotating on the left side of the helicopter 07, in which the airspeed V is reduced. Realization of power device.

That is, in order to maintain the horizontal balance during the forward flight of the helicopter 07, as described above, the angle of attack of the blade 06 rotating on the right side of the helicopter 07 where the airspeed V increases, as described above. As compared with α, the attack angle α of the blade 06 rotating on the left side of the helicopter 07 whose airspeed V decreases is increased, so that the lift L generated on the left and right blades 06 of the helicopter 07 becomes uniform. However, depending on the increase in the payload or the increase in the flight speed, horizontal flight may not be maintained unless the elevation angle α of the left blade 06 is set to a larger elevation angle α.

[0025] For this, the magnitude of lift L generated in a helicopter blade 06 to the size of the angle of attack alpha, so-called, is a change with respect to change in the angle of attack alpha coefficient of lift C _l that dimensionless lift L, In addition to increasing the magnitude of the lift inclination, the angle of attack α, which suddenly decreases after the lift L that increases with the increase in the angle of attack α reaches a maximum value, so-called stall angle, is increased. Thus, the realization of a high-lift device for increasing the maximum lift coefficient C _lmax of the airfoil 01 is required to improve forward flight performance and payload.

Furthermore, in order to reduce the required horsepower at the time of hovering specific to the helicopter 07 and to improve the hovering performance, which flies at a substantially constant position in the air, the airfoil 01 is also required to realize the following performance. You.

(3) Realization of an increase in the lift-drag ratio C ₁ / C _d , which is the ratio between the lift coefficient C _l and the resistance coefficient C _d when hovering, evaluated at the lift coefficient C _l = 0.6. Etc. are required.

However, in response to these requirements, the conventional airfoil 01 has a shape close to a symmetrical airfoil as described above, and as shown in the airfoil data of Table 4, the pitching moment M Although the dimensionless pitching moment coefficient C _m can be reduced, the maximum lift coefficient C
_lmax is small, it is difficult to generate a large lift, and the resistance divergence Mach number M _dd and the lift-drag ratio C _l / C _d are relatively small. As described above, the maximum take-off weight increases and the flight speed increases. There has been a problem with the application of the helicopter 07 to the blade 06 in recent years, which has been demanded to be modified.

[0029]

The helicopter blade airfoil of the present invention has a smaller pitching moment coefficient, a higher resistance divergence Mach number and a higher maximum lift coefficient than the conventional blade airfoil. And a wing type that can achieve a high lift-to-drag ratio, can respond to the increase in maximum take-off weight at the time of helicopter takeoff, which has been progressing rapidly in recent years, and the increase of the flight speed at the time of forward flight. An object of the present invention is to provide a blade airfoil capable of improving hovering performance in hovering and the like.

[0030]

For this reason, the blade airfoil according to claim 1 has the following means (1). (1) The blade thickness at each position in each chord direction of the airfoil is calculated by the aforementioned equation (1), and the coefficients a ₀ ,
a ₁ , a ₂ , a ₃ , a ₄ , a _5, and a ₆ are to be airfoiled in a shape calculated using a numerical value in the range of ± 3% of the numerical value shown in Table 1. did.

That is, the blade airfoil according to the first aspect of the present invention delays stall and generates a high maximum lift. Therefore, as shown by a thick line A in FIG. 1, the blade airfoil is larger than the conventional airfoil. It has a leading edge radius and has a large camber shown by a thick line in FIG. 1 from a shape close to the conventional symmetric airfoil shown by a thin line in FIG. 5 or FIG. With a large asymmetrical wing.

In order to suppress an increase in resistance during high-speed flight and to realize a high-resistance divergence Mach number, the central portions B and C of the upper and lower surfaces of the airfoil are replaced with the conventional airfoil. The one having a radius of curvature larger than the central arc surface, in other words, a shape close to a flat shape was used.

Further, the blade airfoil, particularly the blade airfoil used as the rotor blade of the helicopter, is important, and the pitching moment coefficient, which has been nearly zero, has been reduced to the same size as the conventional one. In order to suppress the airfoil, the shape near the trailing edge of the airfoil was formed to be a convex shape having a so-called reverse camber property having a downwardly convex shape as shown by a portion D.

(A) Thus, in the blade airfoil of the present invention, by using an airfoil having a large leading edge radius, rapid changes in pressure and speed at the blade leading edge can be suppressed.
By causing separation occurring at the wing leading edge A and causing it to occur on the wake side of the blade surface as much as possible, stall caused by occurrence of separation is delayed so as to occur at a higher angle of attack, It is possible to generate a large lift which becomes a high maximum lift.

Further, by providing a large camber and making the airfoil asymmetrical, a high lift can be generated at the same angle of attack as compared with a conventional airfoil that is close to a symmetric airfoil. As shown in Table 5, a high maximum lift coefficient C _lmax with a high maximum lift coefficient of 1.47 should be used. become able to.

[0036]

[Table 5]

Further, by increasing the camber, a high lift can be generated even when the thickness of the airfoil is reduced. Further, by reducing the thickness of the airfoil compared with the conventional airfoil, The resistance of the airfoil during hovering can be reduced to a smaller value of about 0.0085, thereby realizing a high lift-drag ratio with a lift-drag ratio of 70.5 and suppressing an increase in resistance during high-speed flight. In addition, a high resistance divergence Mach number can be realized, and an increase in resistance during low speed flight can be suppressed.

Further, by flattening the B portion on the upper surface and the C portion on the lower surface of the airfoil, it is possible to suppress the generation of a shock wave at a high speed until the speed becomes closer to sonic flight, and to reduce the resistance at a high speed. And an airfoil capable of realizing a high resistance divergence Mach number capable of improving the resistance divergence Mach number to about 0.82.

Further, by providing a cam portion in the direction opposite to the front of the airfoil by providing a cam portion in the vicinity of the trailing edge of the airfoil, as indicated by a portion D having a downward convex shape, the trailing edge of the airfoil is provided. In the vicinity, a downward lift is generated near the attack angle of 0 °, and the lift generation point is separated from the aerodynamic center CP.
The moment arm can generate a pitching moment in the opposite direction near the trailing edge of the airfoil, which cancels the pitching moment generated in the entire airfoil except for the trailing edge near the attack angle of 0 °. Despite the provision of a large camber, the pitching moment coefficient of the entire airfoil can be reduced to 0.03 or less, and can be reduced as in a conventional symmetric airfoil.

Therefore, the invention of claim 1 can be used more effectively when used on the inner wing side of the blade.

The blade airfoil according to claim 2 of the present invention has the following means (2) in addition to the above-mentioned means (1).

(2) The coefficients a ₀ , a ₁ ,
The values of a ₂ , a ₃ , a ₄ , a ₅ and a _{6 used} in Table 1 are different from those of Table 1 in that the invention of claim 1 uses a value within ± 3% of the value described in Table 1. ± 3 of the numerical value described in
The value within the% range was used.

That is, in the blade airfoil of the present invention, an asymmetric airfoil is formed in the same manner as the blade airfoil of the first aspect of the present invention, and the upper and lower surfaces of the airfoil are made flat and the trailing edge is formed. The shape in the vicinity was made to be convex downward.

As a result, the same operation and effect as the above-mentioned (a) can be obtained. In addition, in the blade airfoil of the present invention, the maximum lift coefficient is 1 in comparison with that of the first aspect of the present invention. .32 and the resistance coefficient is slightly increased to 0.0102, but the resistance divergence Mach number is 0.85, which is a large high resistance divergence Mach number. An increase in resistance can be prevented.

Further, the pitching moment coefficient of the entire airfoil can be reduced to 0.01 or less as compared with that of the first aspect of the present invention, and the structural load applied to the pitch angle control device can be reduced. A blade that can be made smaller, has a higher speed compared to the inner wing side, and has a larger moment arm required for pitch angle control.It can be used more effectively on the outer wing side of the blade. Becomes

[0046]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS Hereinafter, a first embodiment of a blade airfoil according to the present invention will be described with reference to the drawings. FIG. 2 is a diagram showing a first embodiment of the blade airfoil of the present invention, using coefficients a ₀ , a ₁ , a ₂ , a ₃ , a ₄ , a ₅ and a ₆ shown in Table 1. This is an airfoil 1 calculated by the equation of Equation 1. Table 6 shows airfoil coordinates that substantially define the airfoil 1. If the values of the coefficients a ₀ , a ₁ , a ₂ , a ₃ , a ₄ , a ₅ and a ₆ are within ± 3% of the values shown in Table 1, the above airfoil 1 Becomes substantially the same airfoil.

[0047]

[Table 6]

In the airfoil 1 of the first embodiment shown in FIG. 2, the airfoil 1 is used on the root side of the 80% span position of the blade 06 of the helicopter 07 to obtain a suitable airfoil 1. The blade thickness ratio t of 1 is as large as 9%.

As described above, when the blade 06 rotates in the direction opposite to the traveling direction, the airspeed V seen from the blade 06 is calculated from the reverse speed component of the rotation speed R by the helicopter 07.
Of the forward speed F of
In order to maintain a balance with the blade 06, which turns in the traveling direction and the airspeed V increases, the angle of attack α is increased. However, since the speed is further reduced inside the blade 06, even if the angle of attack α is large, The maximum lift coefficient C _lmax must be increased so that a large lift L is obtained without stalling.

For this reason, as shown in the figure, the blade airfoil of the present embodiment delays stall so as not to occur until a high angle of attack occurs, and generates a high maximum lift. As shown in the figure, in order to generate a large leading edge radius and to generate a high maximum lift, the center of the blade thickness at each station X / C in the chord direction is connected from a shape close to the conventional symmetrical airfoil. Da camber line, leading edge 04 and trailing edge 05
An asymmetrical airfoil having a large camber, which is greatly shifted upward from the chord line CL connecting the straight line and.

Further, in order to realize a high-resistance divergence Mach number for suppressing an increase in the resistance D during high-speed flight, the airfoil 1
The shapes of the central portions B and C of the upper surface 02 and the lower surface 03 of each of the conventional airfoils having a small curvature are made to have a shape close to a flat shape having a larger radius of curvature than the arc surface of the central portion of the conventional airfoil.

[0052] The pitching moment coefficient C _m, order to suppress the conventional airfoil 01 similar size, a trailing edge 05 near the shape was that of the reverse camber shape which is convex downward shape as shown in part D . Although not shown in the drawing, the end of the trailing edge 05 formed in an inverted camber shape is formed with a short vertical plane because the upper surface 02 and the lower surface 03 are formed by bonding together.

As a result, in the blade airfoil of the present embodiment, since the airfoil 1 is provided with a large leading edge radius, rapid changes in pressure and speed at the leading edge A of the blade can be suppressed. The airflow flowing along the leading edge A of the wing flows along the wing surface further downstream, so that separation can be caused on the wake side of the blade surface as much as possible. Can be delayed so that a high maximum lift can be obtained.

Also, a large camber projecting upward is provided in almost all portions of the airfoil 1 except for the trailing edge 05, so that the airfoil 1 has a strong asymmetry. At the angle α, a high lift can be generated as compared with the symmetric airfoil, and further, the lift at an elevation angle of 0 can be increased, so that a high maximum lift coefficient can be realized.

Further, by increasing the size of the camber, a high lift can be generated even if the thickness of the airfoil 1 is reduced, and as described above, a high maximum lift coefficient can be obtained. Since the thickness of the airfoil 1 is smaller than that of the conventional airfoil 01, the resistance D can be reduced.

As a result, a high lift-drag ratio can be realized, the resistance D at the time of high-speed flight can be suppressed from increasing, and the resistance divergence Mach number _{Mdd at} which the resistance increases sharply becomes closer to the speed of sound. The airfoil 1 having a high resistance divergence Mach number can be realized.

The lift / drag ratio of the so-called inner wing used on the root side as in the airfoil 1 of the first embodiment is to be evaluated at a Mach number M = 0.5. ,
The resistance coefficient C _d and the lift-drag ratio C ₁ / C _d of the airfoil data of the first embodiment shown in Table 5 indicate data of M = 0.5.

Also, by flattening the B portion on the upper surface and the C portion on the lower surface of the airfoil 1, the generation of shock waves at high speed can be reduced.
Around the Mach number close to the speed of sound, it is possible to reduce the resistance D during high-speed flight, in combination with the thinner airfoil 1, and to realize a high-resistance divergence Mach number. Flight up to a nearby Mach number M becomes possible.

In the vicinity of the trailing edge of the airfoil 1, unlike the other parts of the airfoil 1, the camber line is shifted downward from the chord line CL by a large amount and has a downwardly convex D portion. By providing a reverse camber of the shape shown, near the trailing edge of the airfoil,
Near the angle of attack 0 °, a downward lift L is generated, and the position where the downward lift L is generated is determined by the aerodynamic center CP.
(X / C = 0.25), the pitching moment M in the opposite direction is generated near the attack angle of 0 °, which counteracts the pitching moment M generated in the entire airfoil except for the trailing edge 05. In spite of the fact that the airfoil 1 is changed from a symmetric airfoil to an asymmetrical airfoil provided with a large camber, the pitching moment coefficient C of the entire airfoil 1 is increased.
_m can be as small as a conventional symmetric wing.

That is, in the blade airfoil of this embodiment, as shown in Table 5, the maximum lift coefficient C _lmax can be _increased as compared with the conventional airfoil 01 shown in Table 4. The resistance coefficient C _d when the lift coefficient C _l during hovering becomes 0.6 can be reduced, and the lift-drag ratio C _l / C _d can be increased. Further, the resistance divergence Mach number _Mdd can be reduced as compared with the conventional airfoil 01.

In Table 4, M of NACA0006
Compared to _dd , M _dd in the present embodiment is small. This is because the blade thickness ratio t is 6% in the former case, whereas the data shown in Table 5 is 9%. It can be seen that when compared with the same blade thickness ratio t, the high resistance divergence Mach number can be obtained by NACA0012 and NACA
It can be easily estimated from the data of 23012.

[0062] Also, among the blades for airfoil, is intended to be used in the inner airfoil, such performance, it is necessary to improve without increasing the pitching moment coefficient C _m, used in the inner blade side The airfoil that is used is 0.0
A pitching moment coefficient of 3 or less is a general value, and the airfoil of the present embodiment satisfies any Mach number M.

FIG. 4 shows the values of the maximum lift coefficient C _lmax and the resistance divergence Mach number M _dd of the airfoil 1 of this embodiment in NA.
This shows a comparison with other airfoils of the conventional blade airfoil including the CA00XX system. In this figure, the higher the performance is at the upper right of the figure, the higher the performance is. It can also be seen from this figure that the helicopter blade airfoil in the form is superior to the conventional one.

Next, FIG. 3 is a view showing a second embodiment of the blade airfoil according to the present invention, in which coefficients a ₀ ,
Using a ₁ , a ₂ , a ₃ , a ₄ , a ₅ and a ₆ ,
Is the airfoil 2 calculated by the following equation. Table 7 shows airfoil coordinates that substantially define the airfoil 2. Note that the coefficients a ₀ ,
The values of a ₁ , a ₂ , a ₃ , a ₄ , a ₅ and a ₆ are shown in Table 2
If the value is within ± 3% of the value shown in the above, the airfoil is substantially the same as the airfoil 2 described above.

[0065]

[Table 7]

In the airfoil 2 of the second embodiment shown in FIG. 3, the blade is used on the blade tip side of the blade 06 of the helicopter 07 with respect to the 80% span position. Is as small as 7%.

When the blade 06 rotates in the same direction as the traveling direction, the airspeed V viewed from the blade 06 is the sum of the forward speed component of the rotation speed R and the forward speed of the helicopter 07. The high speed region is on the wing tip side where the blade airfoil of the form (1) is used.

For this reason, it is required that the airfoil 2 has a large resistance divergence Mach number _Mdd so that an increase in resistance is suppressed even in a high speed range. Therefore, as described above, the airfoil 1 of the first embodiment has a particularly large maximum lift coefficient C _lmax at low speed, and also has a resistance divergence Mach number M _dd .
Whereas was made in relatively high airfoil 1, the airfoil 2 of the present embodiment, in particular, drag divergence Mach number M _dd is large and, for the maximum lift coefficient C _lmax at low speed,
It is necessary to have a relatively high airfoil 2.

For this reason, as shown in the figure, also in the blade airfoil of the present embodiment, the first embodiment shown in FIG.
As in the case of the blade airfoil of the form, an asymmetric airfoil having a camber projecting upward at most of the stations X / C in the chord direction is provided, and the upper surface 02 and the lower surface 03 of the airfoil 2 are provided. Was flattened, and the shape near the trailing edge 05 was provided with an inverted camber that became convex downward.

As a result, the same functions and effects as those of the airfoil 1 of the first embodiment described above can be obtained, and the blade airfoil of the present embodiment can be compared with the airfoil of the first embodiment. Then, the maximum lift coefficient C _lmax is slightly reduced to 1.32, and the resistance coefficient C _d is also slightly increased.
The resistance divergence Mach number M _dd becomes larger, so that it is possible to prevent a sudden increase in resistance until a higher-speed flight, and to make the pitching moment coefficient C _m of the entire airfoil 2 smaller. The load on the outer wing of the helicopter blade 06 can be reduced because the load on the structure applied to the pitch angle control device can be reduced, the speed becomes higher than that on the inner wing side, and the moment arm required for pitch angle control becomes larger. It can be more effective when used.

That is, as shown in Table 5, the blade airfoil according to the present embodiment has a maximum lift coefficient C _lmax and a lift-drag ratio C ₁ / C, as compared with the first embodiment shown in the table. Although the increase in _d is small, the increase in the resistance divergence Mach number M _dd is remarkable, and the pitching moment coefficient C _m should be approximately the same as that of the conventional airfoil 01 shown in Table 4. Can be.

The lift / drag ratio C ₁ / C of the outer wing used on the wing tip side as in the airfoil 2 of the second embodiment.
_{Since d} is to be rated at Mach number M = 0.6, the drag coefficient and lift / drag ratio C _l / C _d of the second form airfoil data shown in Table 5 are as follows: This shows data of 0.6.

As described above, in the airfoil 2 according to the second embodiment, as shown in FIG. 4, the airfoil has a large resistance divergence Mach number _Mdd , and includes the airfoil 1 according to the first embodiment. Also comes to the upper right side of the resistance divergence Mach number _Mdd region of the airfoil, which is said to be advanced, and it can be seen that the airfoil can have a high performance.

Further, among blade airfoils used on the wing tip side, it is necessary to improve such performance without increasing the absolute value of the pitching moment M. A typical value is 0.01 or less for the airfoil that is used, and the airfoil 2 of the present embodiment satisfies both cases.

As described above, if the blade airfoil of the first embodiment is used as an inner airfoil and the blade airfoil of the second embodiment is used as an outer airfoil, the pitching moment coefficient While maintaining a small airfoil, the airfoil can achieve a high resistance divergence Mach number, a high maximum lift coefficient and a high lift-drag ratio, further increasing the maximum takeoff weight during helicopter takeoff, and enabling faster forward flight. Furthermore, helicopter 0 that can improve hovering performance
7 can be realized.

[0076]

As described above, in the blade airfoil of the present invention, the blade thickness on the upper surface side and the lower surface side at each chord direction position of the airfoil is expressed by the following equation. The coefficients a ₀ , a ₁ , a ₂ , a ₃ , a ₄ , a _5, and a _{6 in} the equation (1) are set to numerical values in the range of ± 3% of the numerical values shown in Table 1. .

Accordingly, the blade airfoil according to the present invention has a large camber except for a large leading edge radius and a trailing edge shape.
It has a flat upper and lower surface and a downwardly convex trailing edge shape,
It can have high maximum lift, high resistance divergence Mach number, high lift-drag ratio and low pitching moment, making it suitable for use on the inner wing side of a helicopter blade.

In the blade airfoil of the present invention,
Is a coefficient a in the equation (2).₀, A₁, A _Two, A_Three, A_Four,
a_FiveAnd a₆Is a value within ± 3% of the value shown in Table 2.
It is what has been.

As a result, a higher resistance divergence Mach number can be obtained, and the pitching moment can be made extremely small, close to zero, as in the case of the conventional blade airfoil. It can be made suitable for.

[Brief description of the drawings]

FIG. 1 is a sectional view of an airfoil showing a comparison between a blade airfoil of the present invention and a conventional blade airfoil.

FIG. 2 is an airfoil sectional view showing a first embodiment of a blade airfoil according to the present invention.

FIG. 3 is an airfoil sectional view showing a second embodiment of the blade airfoil of the present invention.

FIG. 4 is a diagram showing the relationship between the maximum lift coefficient and the resistance divergence Mach number of the conventional blade airfoil and the blade airfoil according to the first and second embodiments.

FIG. 5 shows a conventional blade airfoil NACA00X.
X type airfoil cross section

FIG. 6 is a diagram showing pneumatic force acting on a blade airfoil.

FIG. 7 is a diagram showing the operation of a blade during forward flight of a helicopter.

[Explanation of symbols]

01 ... airfoil 02 ... upper surface 03 ... lower surface 04 ... front edge 05 ... trailing edge 06 ... blade 07 ... helicopter 1 ... airfoil (first embodiment) 2 ... airfoil (second embodiment) a _0, a _{1, a 2, a 3,} a 4, a 5, a 6 ... coefficient y ... blade thickness direction coordinate x ... chordwise coordinate C ... chord length CL ... chord line t ... TsubasaAtsuhi C _l ... lift coefficient C _lmax ... maximum lift coefficient C _d ... resistance coefficient C _m ... pitching moment coefficient CP ... aerodynamic center F ... forward speed R ... rotational speed V ... airspeed alpha ... attack angle M ... Mach number M _dd ... drag divergence Mach number L ... Lift D ... Drag (resistance) M ... Pitching moment ρ ... Air density A ... Atmospheric flow direction

Claims (2)

[Claims] 1. The blade thickness at each chordwise position in each span direction of the airfoil is represented by the following equation (1), and the coefficients a ₀ , a ₁ , a _{2, a 3, a 4,} a 5 and a ₆ are blade for airfoil, characterized in that it is a number in the range of ± 3% of the numerical values shown in Table 1. (Equation 1) Here, y indicates a blade thickness direction coordinate, x indicates a chord direction coordinate, c indicates a chord length, and t indicates a blade thickness ratio (%), which is a ratio between the maximum blade thickness of the airfoil and the chord length c. Shall be. [Table 1] 2. The blade thickness at each chord direction position in each span direction of the airfoil is represented by the following equation (2), and the coefficients a ₀ , a ₁ , a _{2, a 3, a 4,} a 5 and a ₆ are blade for airfoil, characterized in that it is a number in the range of ± 3% of the numerical values shown in Table 2. (Equation 2) Here, y indicates a blade thickness direction coordinate, x indicates a chord direction coordinate, c indicates a chord length, and t indicates a blade thickness ratio (%), which is a ratio between the maximum blade thickness of the airfoil and the chord length c. Shall be. [Table 2]