JP2018131038A - Rotor, drone and helicopter - Google Patents

Abstract

[PROBLEMS] To secure high torsional rigidity while obtaining a high figure of merit. A rotor 6 having a solidity of 10% or less and two blades 3 are provided. The shape of each blade 3 has a maximum chord length in an area with a radius of 10% to 35%, the chord length decreases from the position of the maximum chord length to the tip direction, and the reduction of the chord length decreases at a predetermined position. The chord length decreases gradually from the position toward the tip, and the inflection point is in the region of radius 70% to 95%, the tip is elliptical, and touches the chord at two points. The maximum chord length is more than three times the chord length of the imaginary wing tip created by the tangent at the wing tip, and the constriction amount in the middle of the radius is less than 30% of the imaginary wing tip. The angle becomes maximum near the maximum chord length, the chord length decreases from the maximum to the tip, and the mounting angle of the root decreases from the position where the mounting angle reaches the maximum to the root. Yes. [Selection] Figure 1

Description

  The present invention relates to a rotor used in a drone, a helicopter, and the like, and a drone and a helicopter having such a rotor.

  In recent years, a drone called a drone having four or more rotors has been actively used for photographing and observation purposes. A drone normally uses a plastic two-bladed fixed-pitch rotor, and the performance of the rotor affects the weight of the payload and the flight time. The rotor shape varies, and it seems that the classic propeller theory (Non-patent Document 1) is applied from a simple taper. The chord length of the root is wide, and then the chord length is sharply reduced from there to the tip. The taper ratio shifts to a shape with a rounded tip. Further, GEN H? 4 (Non-Patent Document 2) developed by GEN CORPORATION as a manned small helicopter similarly uses a fixed pitch rotor, and adopts a simple taper for the rotor.

  In the field of helicopters, Westland Linx in the UK has achieved world speed records by using a rotor with a wider swivel length called BERP (Non-patent Document 2, Patent Document 1) and a receding angle. . In this rotor, the discontinuous leading edge with the enlarged chord at the tip functions as a dog tooth used in airplanes, generating vertical vortices to suppress the separation of the upper surface and earning a lift coefficient on the reverse side, and at the tip The purpose is to reduce the wave resistance on the forward side by the receding angle. The helicopter must trade off the actual lift with the theoretical maximum lift required from the simple momentum theory, rotor rotation area, and input power, and the maximum speed must be traded off, so the rotor shape is rectangular. Many of them have a shape with a weak taper at the tip. On the other hand, since the drone takes a long time to hover, there is a high demand for using a rotor with a high figure of merit to increase the flight time and payload.

US Pat. No. 5,174,721

Adkins, C. N., Liebeck, R. H., "Design of Optimum Propellers," AIAA-83-0190, Jan. 1983 Yanagisawa Gennai, counter-rotating single-seat 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

  As an effective method for designing a rotor with a large figure of merit, there is a method by Adkins et al. However, the method of Adkins et al. Could not reflect the contracted flow (reduction in the radius of the cylindrical flow) or the vortex entrainment (roll-up) seen in the actual wake of the rotor. Therefore, the figure of merit was not the maximum. In the method of Adkins et al., Since the chord length is too large near the root, it is necessary to correct the chord length near the blade root to a certain value or less. However, the mounting angle is not corrected at this time. It was not the optimal mounting angle when the chord length was limited. Since the rotor used for the drone is generally small, the Reynolds number of the airfoil is low. In order to generate a large lift coefficient at this low Reynolds number, an airfoil with a thin and large camber is used. Since it has a large camber, the moment coefficient Cm takes a large negative value. For this reason, a moment to twist the blade works, but the blade obtained by the method of Adkins et al. Has a constricted middle part of the blade radius and uses a thin airfoil as described above, so it has rigidity to withstand twisting. There is a problem of being twisted and twisting.

  In view of the circumstances as described above, 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 helicopter having such a rotor.

In order to achieve the above object, a rotor according to an aspect of the present invention includes:
A rotor with a solidity of 10% or less,
With two blades,
The shape of each blade is
Has a maximum chord length in the range of 10% to 35% radius,
The chord length decreases from the position of the maximum chord length to the tip direction,
The reduction of the chord length becomes maximum at a predetermined position, and the reduction of the chord length becomes gentle from the position toward the tip,
There is an inflection point in the area of radius 70% to 95%,
The tip is oval,
The maximum chord length is more than three times the hypothetical chord length of the hypothetical tip created by the tangent that touches the chord at two points.
The amount of constriction in the middle of the radius is 30% or less of the virtual wing tip,
The mounting angle is maximum near the maximum chord length, the chord length decreases from the maximum position to the tip,
The attachment angle of the base decreases from the position where the attachment angle takes the maximum value to the base.

  The shape of each of the blades is preferably a shape in which the chord length is a curve having a first-order differential from the root to the tip.

  The drone which concerns on another form of this invention has the rotor of the said structure.

  A helicopter according to still another embodiment of the present invention has the rotor configured as described above.

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

It is a top view of the rotor which concerns on one Embodiment of this invention. The trajectory of the rotor tip vortex is shown. a) is a locus of the rotor blade tip vortex by the method of Adkins et al., and b) is a locus of the rotor blade tip vortex by the Harada vortex method of the present inventor. a) is the actual tip vortex rollup, and b) is the modeled tip vortex rollup. It is a graph which shows the optimal chord length distribution of a conventional method. It is a graph which shows the optimal attachment angle of the conventional method. It is a graph which shows the optimal chord length distribution of the conventional method which restricted the chord length. It is a graph which shows the optimal attachment angle of the conventional method which restricted 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 characteristic of the attachment angle of this invention. It is a figure for demonstrating Harada's vortex method, and has shown the coordinate system and discharge | emission vortex of a propeller. It is a figure for demonstrating Harada's vortex method, and has shown the enlarged view of the 1st blade in FIG. It is a figure for demonstrating Harada's vortex method, and has shown the cross section and inflow velocity of a blade. It is a figure for demonstrating Harada's vortex method, and has shown the force which acts on a 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 includes two blades 3 and a hub 4 having a shaft hole 5.

Here, with the method of Adkins et al., As shown in FIG. On the other hand, the shape of the wake can be further determined as follows by using Harada's vortex method (see Non-Patent Document 5, which will be described later) as the present inventor.
1) As shown in FIG. 2 b), the tip vortex radius r is expressed by the following equation: r / R = A + (1−A) e− ^λψ (1)
Shrink according to Here, R is the rotor radius, and A is a constant obtained from experiments, and 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 discharge vortices in the range of the pitch of the spiral vortex from the blade tip are immediately rolled up to become the blade tip vortex 1 and move as a group.
3) The tip vortex 1 moves at a speed about half the moving speed of the vortex inside the tip vortex 1.
4) The vortex inside the rotor rotating surface 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 discharge vortex shrinks, and moves at the speed of v ′ given by the following equation.

v ′ = v (R / r) ² (2)
5) Since simulation of the blade tip vortex 1 is difficult, as shown in FIG. 3 b, a model 2 is used in which a group of vortex vortices are moved at a constant speed without moving in the radial direction.

  By using such a wake vortex and minimizing the chord length near the root that is too large by the method of Adkins et al., The evaluation function obtained from Harada's vortex method under the constraint condition of the chord length, A rotor 6 having an appropriate chord length can be obtained.

In this embodiment, the rotor 6 having the following shape is obtained.
1. 1. a rotor 6 having a solidity of 0.1 or less; The number of blades 3 is 2,
3. The chord length is a curve with a continuous first derivative from the root to the tip.
4). Has a maximum chord length from 10% to 35% radius,
5. 5. From there, the chord length decreases toward the tip. Once the decrease reaches its maximum, the decrease becomes gradual, and there is an inflection point from radius 70% to 95%.
7). The tip is oval,
8). The maximum chord length is more than three times the hypothetical chord length of the hypothetical tip created by the tangent that touches the chord at two points.
9. The amount of constriction in the middle of the radius is 30% or less of the virtual wing tip,
10. 10. The mounting angle is 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 reached to the base.

  By using the rotor 6 of this shape, high torsional rigidity can be secured 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. 2 a), and FIG. 5 shows the mounting angle distribution. Here, the calculation conditions are shown in Table 1.

  This shape is not practical because the root chord length reaches 12 cm as shown in FIG. Among commercially available drones, there is an example in which the chord length at the base of the blade 3 is simply cut to a practical length.

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

  The value at the wing tip C of the line segment AB in contact with the obtained chord distribution at two points, the imaginary wing tip chord length D is about 1.36 cm. On the other hand, at 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. In addition, the chord length gradually decreases in the vicinity of the wing tip, but 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.

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

  FIG. 8 shows the chord length distribution and FIG. 9 shows the installation angle distribution when the Harada's vortex method uses the model of FIG. The figure of merit at this time is 89.5%.

  A line segment AB in contact with the obtained chord distribution at two points is a value at the blade tip C, and a virtual blade tip chord length D is about 1.25 cm. On the other hand, at 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. In addition, the chord length gradually decreases in the vicinity of the blade tip, but suddenly decreases in the vicinity of the blade tip C, and the blade tip has an elliptical shape. The amount of constriction appearing in the middle of the radius indicated by M is 0.14 cm, which is suppressed to 11% of the wing tip chord length.

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

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

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

  The present invention is not limited to the above-described embodiment, and various modifications are possible within the scope of the technical idea, and the scope 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 propeller coordinate system and the discharge vortex. It is considered that the propeller moves in the x-axis direction while rotating, and a discharge vortex is left in the moved locus.

FIG. 11 shows an enlarged view of the first blade. There are i-th representative point from the axis of rotation on r _i apart blades, j-th emission vortices are discretized into short line segments as shown in Fig white circles. Assuming that the induced velocity in the x direction caused by the discharge vortex of the j-th unit intensity at the i-th representative point is X _ij and the induced velocity in the z-direction is Z _ij , the induced velocity u in the x-direction caused by the i-th representative point u. _{The i} and z-direction guiding speeds w _i are respectively 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 Biosavart's law, and г _j is the j-th emission vortex.

FIG. 12 shows the cross section of the blade and the inflow speed. 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. Here, U is the propeller forward speed. The inflow angle φ _i and the inflow speed V _i are given by the following equations, respectively.

^{_{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 blade element is given by the Kutta-Jukovsky theorem as follows: 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, 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)
The local resistance dD _i acting on the i th blade element is given by the following equation using the resistance coefficient C _D.

_{dD i = 1/2 · ρV} i 2 C 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 .
_{dT i = dL i cosφ i -dD} i sinφ i (13)
dN _i = dL _i sin φ _i + dD _i cos φ _i (14)
It becomes. Since the local absorption power dP _i is dN _i r _i Ω, it is given by the following equation.

dP _i = (dL _i sin φ _i + dD _i cos φ _i ) r _i Ω (15)
Eventually, the thrust and absorption power are calculated from the equations (13) and (15) as follows: T = BΣdT _i (16)
P = BΣdP _i (17)
Given in. Here, B is the number of blades.

The optimal design problem for the propeller is C _L , C _D , Ω, design power P ₀ and гi as unknown

[Condition]: P = P ₀
Under

[Objective function]: -T
This results in an optimization problem that minimizes.

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

Claims (4)

A rotor with a solidity of 10% or less,
With two blades,
The shape of each blade is
Has a maximum chord length in the range of 10% to 35% radius,
The chord length decreases from the position of the maximum chord length to the tip direction,
The reduction of the chord length becomes maximum at a predetermined position, and the reduction of the chord length becomes gentle from the position toward the tip,
There is an inflection point in the area of radius 70% to 95%,
The tip is oval,
The maximum chord length is more than three times the hypothetical chord length of the hypothetical tip created by the tangent that touches the chord at two points.
The amount of constriction in the middle of the radius is 30% or less of the virtual wing tip,
The mounting angle is maximum near the maximum chord length, 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 to the base. The rotor according to claim 1,
The shape of each blade is
The chord length is a curve with a continuous first-order differential from the root to the tip.   A drone having the rotor according to claim 1.   A helicopter having the rotor according to claim 1 or 2.