JP6856930B2 - Rotor, drone and helicopter - Google Patents

Description

The present invention relates to rotors used for drones, helicopters, etc., and drones and helicopters having such rotors.

In recent years, flying objects called drones with four or more rotors have been actively used for photography and observation purposes. Drones usually use a two-bladed fixed-pitch rotor made of plastic, and the performance of this rotor affects the weight of the payload and the flight time. There are various rotor shapes, and it seems that the classical propeller theory (Non-Patent Document 1) was applied from the simple taper. Some of them have a taper ratio and a rounded tip. Further, GEN H-4 (Non-Patent Document 2) developed by GEN CORPORATION as a manned small helicopter also uses a fixed pitch rotor, and adopts a simple taper for the rotor.

In the field of helicopters, Westland Linx of the United Kingdom has achieved a world speed record by using a rotor with an extended tip chord length called BERP (Non-Patent Document 2 and Patent Document 1) and a receding angle. .. This rotor functions as a dog tooth used in an airplane with a discontinuous front edge that expands the wing chord at the tip, and generates a vertical vortex to suppress peeling of the upper surface, thereby gaining a lift coefficient on the backward side and the tip. The purpose is to reduce the wave-making resistance on the forward side by the receding angle of. Since the helicopter must trade off the maximum speed with the figure of merit (Non-Patent Document 3) in which the actual lift is divided by the theoretical maximum lift obtained from the simple momentum theory, the rotor rotation area, and the input power, the rotor shape is rectangular. Most of them have a shape with a weak taper at the tip. On the other hand, since drones have a long hovering time, there is a high demand for using a rotor with a high figure of merit to increase flight time and payload.

U.S. Pat. No. 5,174,721

Adkins, C.N., Leebeck, R.H., "Design of Optimum Propellers," AIAA-83-0190, Jan. 1983. Gennai Yanagisawa, Contra-rotating single-seater helicopter GEN H-4 Development and future prospects, Heli Japan 2002, November 2002 https://en.wikipedia.org/wiki/BERP_rotor R. W. Prouty, "Helicopter Performance, Stability, and Control," 1990 Masashi Harada, "Design method for low Reynolds number propellers," JAXA-RR-06-032, 2007

There is a method by Adkins et al. (Non-Patent Document 4) as an effective method when designing a rotor having a large figure of merit. However, the method of Adkins et al. Could not reflect the contraction (reduction of the radius of the tubular flow) and the vortex entrainment (roll-up) seen in the wake of the actual rotor. Therefore, the figure of merit was not the maximum. In addition, since the chord length near the root of the blade becomes too large in the method of Adkins et al., It is necessary to correct the chord length near the root of the blade to a certain value or less, but the mounting angle is not corrected at this time. , It was not the optimum mounting angle when the chord length was limited. Since the rotor used for drones is generally small, the airfoil Reynolds number is low. An airfoil with a thin and large camber is used to generate a large lift coefficient at this low Reynolds number. Since it has a large camber, the moment coefficient Cm takes a large negative value. For this reason, a moment that twists the blade works, but the blade obtained by the method of Adkins et al. Has a constricted mid-blade radius and uses a thin airfoil as described above, so it has rigidity to withstand twisting. There is a problem of twisting due to lack of.

In view of the above circumstances, an object of the present invention is to provide a rotor capable of ensuring high torsional rigidity while obtaining a high figure of merit, and a drone and a helicopter having such a rotor.

In order to achieve the above object, the rotor according to one embodiment of the present invention is
A rotor with a solidity of 10% or less,
Has two blades,
The shape of each blade
It has a maximum chord length in the region with a radius of 10% to 35%, and has a maximum chord length.
The chord length decreases from the position of the maximum chord length toward the tip,
The decrease in chord length is maximized at a predetermined position, and the decrease in chord length becomes gradual from that position toward the tip.
There is an inflection point in the area with a radius of 70% to 95%,
The tip is oval,
The maximum chord length is more than three times as long as the hypothetical chord length created by the tangent line that touches the chord at two points.
The amount of constriction in the middle of the radius is less than 30% of the virtual wing tip,
The mounting angle becomes maximum near the maximum chord length, and the chord length decreases from the maximum position to the tip.
The mounting angle of the base decreases from the position where the mounting angle takes the maximum value toward the base.

The shape of each blade is preferably such that the chord length is a curve in which the first derivative is continuous from the root to the tip.

The drone according to another embodiment of the present invention has a rotor having the above configuration.

The helicopter according to still another embodiment of the present invention has a rotor having the above configuration.

According to the present invention, the figure of merit of the rotor is high, and if it is used for a drone or a helicopter, the flight time and the payload weight can be increased. In addition, in the method of Adkins et al., The chord length near the base of the blade became too large, and if this was limited to a certain value or less, the mounting angle was not optimal. According to the present invention, the mounting angle is optimized while limiting the chord length. In the method of Adkins et al., A constriction appears in the chord length in the middle radius, but in the present invention, the chord length in the middle radius can be increased by appropriately limiting the chord length, and the taper ratio can be increased. It can be made larger and have a flat shape that is resistant to twisting.

It is a top view of the rotor which concerns on one Embodiment of this invention. It shows the trajectory of the wing tip vortex of the rotor. a) is the locus of the rotor blade tip vortex by the method of Adkins et al., And b) is the locus of the rotor blade tip vortex by the Harada vortex method of the present inventor. a) is the actual roll-up of the wing tip vortex, and b) is the modeled roll-up of the wing tip vortex. It is a graph which shows the optimum chord length distribution of the conventional method. It is a graph which shows the optimum mounting angle of the conventional method. It is a graph which shows the optimum chord length distribution of the conventional method which limited chord length. It is a graph which shows the optimum mounting angle of the conventional method which limited the chord length. It is a graph which shows the characteristic of the chord length distribution of this invention. It is a graph which shows the feature of the mounting angle of this invention. It is a figure for explaining Harada's vortex method, and shows the coordinate system of the propeller and the emission vortex. It is a figure for demonstrating the vortex method of Harada, and shows the enlarged view of the first blade in FIG. It is a figure for explaining Harada's vortex method, and shows the cross section and the inflow velocity of a blade. It is a figure for explaining Harada's vortex method, and shows the force acting on the blade blade element.

Hereinafter, embodiments of the present invention will be described with reference to the drawings.

FIG. 1 is a top view of a rotor according to an embodiment of the present invention.

As shown in FIG. 1, the rotor 6 has two blades 3 and a hub 4 having a shaft hole 5.

Here, the method of Adkins et al. Can handle only the wake with a constant pitch and a constant radius, as shown in a) of FIG. On the other hand, by using Harada's vortex method (see Non-Patent Document 5, which will be described later), which is the inventor of the present invention, the shape of the wake can be further determined as follows.
1) As shown in b) of FIG. 2, the wing tip vortex radius r is given by the following equation r / R = A + (1-A) e ^−λψ (1).
Shrinks according to. Here, R is the rotor radius, and A is a constant obtained from the experiment, which is 0.76 to 0.8, preferably 0.78. Also, λ is a constant obtained by experiment, and ψ is the blade azimuth.
2) As shown in a) of FIG. 3, the emission vortices in the range of the pitch of the spiral vortex from the wing tip immediately roll up to become the wing tip vortex 1 and move as a group.
3) The wing tip vortex 1 moves at a speed approximately half the moving speed of the vortex inside the wing tip vortex 1.
4) On the rotating surface of the rotor, the internal vortex moves downward at the inflow velocity v derived from the simple momentum theory. According to the above equation (1), the velocity increases as the radius of the emission vortex contracts, and the velocity moves at the velocity of v'given by the following equation.

v'= v (R / r) ² (2)
5) Since it is difficult to simulate the wing tip vortex 1, a model 2 is used in which a group of vortices as a wing tip vortex is moved at a constant velocity without moving in the radial direction as shown in b) of FIG.

By using such a wake vortex and minimizing the chord length near the root, which is too large by the method of Adkins et al., The evaluation function obtained from Harada's vortex method is minimized under the constraint of chord length. It is possible to obtain a rotor 6 having an appropriate chord length.

In this embodiment, a rotor 6 having the following shape is obtained.
1. 1. Rotor 6 with solidity of 0.1 or less and 2. The number of blades 3 is 2,
3. 3. The chord length is a curve with continuous first derivative from the root to the tip.
4. Has a maximum chord length from 10% to 35% radius,
5. 6. The chord length decreases from there to the tip. After the decrease was maximized, the decrease became gradual, with an inflection point from 70% to 95% radius.
7. The tip is oval,
8. The maximum chord length is more than three times as long as the hypothetical chord length created by the tangent line that touches the chord at two points.
9. The amount of constriction in the middle of the radius is less than 30% of the virtual wing tip,
10. The mounting angle becomes maximum near the maximum chord length, and decreases from there to the tip. The attachment angle of the base decreases from the position where the maximum value is taken toward the base.

By using the rotor 6 having this shape, it is possible to secure high torsional rigidity while obtaining a high figure of merit.

FIG. 4 shows the chord length distribution when the Harada vortex method is optimized using the vortex model of FIG. 2a), and FIG. 5 shows the mounting angle distribution. Here, the calculation conditions are shown in Table 1.

This shape is not practical because the chord length at the base reaches 12 cm as shown in FIG. In some commercially available drones, the chord length at the base of the blade 3 is simply cut off to a practical length.

Since Harada's vortex method obtains a solution by minimizing the evaluation function, it is possible to impose a limit on the chord length by adding a penalty function that increases in value when the chord length exceeds 4 cm to the evaluation function. FIG. 6 shows the chord length distribution and FIG. 7 shows the attachment angle distribution when the vortex model of FIG. 2a) is used in the Harada vortex method and the chord length is further limited and optimized. The figure of merit at this time is 89.4%.

The value at the tip C of the line segment AB tangent to the obtained chord distribution at two points, the virtual tip chord length D is about 1.36 cm. On the other hand, in the radius E, the chord length F has a maximum value of 4.79 cm, which is 3.5 times the virtual wing tip chord length D. Further, the chord length gradually decreases near the wing tip, but the shape is cut off at the wing tip C. The constriction in the middle radius indicated by M is 0.32 cm, which is 23% of the chord length of the virtual wing tip.

Further, the obtained mounting angle has a maximum value of 9.32 degrees at a radius I close to a radius E at which the chord length takes a maximum value. The tangent KL to the mounting angle distribution near the base has a positive gradient.

FIG. 8 shows the chord length distribution when the model of b) in FIG. 3 is used in Harada's vortex method, and the chord length is further limited and optimized, and the attachment angle distribution is shown in FIG. The figure of merit at this time is 89.5%.

The line segment AB tangent to the obtained chord distribution at two points is the value at the tip C, and the virtual tip chord length D is about 1.25 cm. On the other hand, in the radius E, the chord length F has a maximum value of 4.94 cm, which is 4.0 times the virtual wing tip chord length D. Further, the chord length gradually decreases near the wing tip, but suddenly decreases near the wing tip C, and the wing tip has an elliptical shape. The amount of constriction that appears in the middle of the radius indicated by M is 0.14 cm, which is suppressed to 11% of the chord length of the virtual wing tip.

Further, the obtained mounting angle has a maximum value of 9.0 degrees at a radius I close to a radius E at which the chord length takes a maximum value. The tangent KL to the mounting angle distribution near the base has a positive gradient.

The rotor 6 designed using the chord length shown in FIG. 9 is shown in FIG.

The rotor 6 according to the present invention can be used for a drone rotor blade, a fixed pitch coaxial counter-rotor type manned helicopter rotor blade, a fixed pitch coaxial double-rotor rotor type unmanned helicopter rotor blade, and the like.

The present invention is not limited to the above-described embodiment, and various modifications can be made within the scope of the technical idea, and the range of the modifications also belongs to the technical scope of the present invention.

Next, the Harada vortex method described above will be described.

FIG. 10 shows the coordinate system of the propeller and the emission vortex. It is considered that the propeller moves in the x-axis direction while rotating, and a emission vortex is left in the moved trajectory.

FIG. 11 shows an enlarged view of the first blade. _{There is an i-th representative point on the blade r i} away from the axis of rotation, and the j-th emission vortex is discretized into short line segments as shown by the white circles in the figure. Assuming that the induction velocity in the x direction caused by the emission vortex of the jth unit intensity at the i-th representative point is X _ij and the induction velocity in the z direction is Z _ij , the induction velocity in the x direction caused by the i-th representative point u _i, each guiding speed _{w i} the z-direction _{u i = ΣX ij г j (} 3)
w _i = ΣZ _ij г _j (4)
Given in. Here, X _ij and Z _ij are constants obtained from Biot-Savart's law, and г _j is the j-th emission vortex.

FIG. 12 shows the cross section of the blade and the inflow velocity. When the tangential component of the air flowing into the blades and _{U T U T} is _{U T = r i Ω-w} i (5)
Given in. Here, Ω is the rotational angular velocity of the propeller. Further, when the axial component of the air flowing into the blades and _{U P U P} is _{U P = U-u i (} 6)
Given in. Where U is the forward speed of the propeller. Each inflow angle phi _i and inlet velocity V _i is given by the following equation.

^{_{φ i = tan -1 (U P}} / U T) (7)
^{_{V i = √ (U P 2}} + U T 2) (8)
FIG. 13 shows the force acting on the blade blade element. _{The local lift dL i} acting on the i-th wing element is based on the Kutta-Jukowski _{theorem dL i = ρV i} г _i db (9).
Given in. Here, ρ is the air density and db is the width of the blade element. Or, by using the lift coefficient _{C L dL i = 1/2} · ρV i 2 C L c i db (10)
It is represented by. Here, c _i is the chord length of the i-th blade element. (9), _{c i} from equation (10) is given by the following equation.

c _i = 2г _i / C _L V _i (11)
i-th local resistance dD _i acting on blade element of is given by the following equation using the resistance coefficient C _D.

dD _i = 1/2 · ρV _i ² _{CD d} c _i db (12)
C _D but is a function of the Reynolds number and C _L, using a constant C _L, C _D is good as a constant as a insensitive to Reynolds number.

The axial component of the resultant force of the local lift dL _i and the local resistance dD _{i is the local thrust dT i} , and the tangential component is dN _i , respectively.
dT _i = dL _i cosφ _{i −} dD _{i sinφ i} (13)
dN _i = dL _i sinφ _i + dD _i cosφ _i (14)
Will be. Since the local absorption power dP _i is dN _i r _i Ω, it is given by the following equation.

dP _i = (dL _i sinφ _i + dD _i cosφ _i ) r _i Ω (15)
After all, the thrust and absorption power are T = BΣdT _i (16) from equations (13) and (15), respectively.
P = BΣdP _i (17)
Given in. Here, B is the number of blades.

Optimal design problem of propeller _C L, _{C D, Ω,} to set the design power _{P 0,} as unknowns гi

[Condition]: P = P ₀
Under

[Objective function]: -T
It comes down to an optimization problem that minimizes.

1 Wing tip vortex 2 Wing tip vortex model 3 Blade 4 Hub 5 Shaft hole 6 Rotor

Claims (4)

A rotor with a solidity of 10% or less,
Has two blades,
The shape of each blade
It has a maximum chord length in the region with a radius of 10% to 35%, and has a maximum chord length.
The chord length decreases from the position of the maximum chord length toward the tip,
The decrease in chord length is maximized at a predetermined position, and the decrease in chord length becomes gradual from that position toward the tip.
There is an inflection point in the area with a radius of 70% to 95%,
The tip is oval,
The maximum chord length is more than three times as long as the hypothetical chord length created by the tangent line that touches the chord at two points.
The amount of constriction in the middle of the radius is less than 30% of the virtual wing tip,
The mounting angle becomes maximum near the maximum chord length, and the chord length decreases from the maximum position to the tip.
The mounting angle of the base decreases from the position where the mounting angle takes the maximum value toward the base. The rotor according to claim 1.
The shape of each blade
The chord length is a rotor with a continuous first derivative from the root to the tip. A drone having a rotor according to claim 1 or 2. A helicopter having a rotor according to claim 1 or 2.

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