WO2017145563A1 - Rotor blade, drone, and helicopter - Google Patents
Rotor blade, drone, and helicopter Download PDFInfo
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- WO2017145563A1 WO2017145563A1 PCT/JP2017/001030 JP2017001030W WO2017145563A1 WO 2017145563 A1 WO2017145563 A1 WO 2017145563A1 JP 2017001030 W JP2017001030 W JP 2017001030W WO 2017145563 A1 WO2017145563 A1 WO 2017145563A1
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- rotor blade
- chord length
- radius
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- 230000007423 decrease Effects 0.000 claims abstract description 3
- 238000000034 method Methods 0.000 description 15
- 238000013461 design Methods 0.000 description 3
- 238000010521 absorption reaction Methods 0.000 description 2
- 238000007796 conventional method Methods 0.000 description 2
- 108010069407 BERP Proteins 0.000 description 1
- 240000005002 Erythronium dens canis Species 0.000 description 1
- 101100126329 Mus musculus Islr2 gene Proteins 0.000 description 1
- 102100038798 Tripartite motif-containing protein 3 Human genes 0.000 description 1
- 230000008602 contraction Effects 0.000 description 1
- 238000011161 development Methods 0.000 description 1
- 238000011156 evaluation Methods 0.000 description 1
- 238000002474 experimental method Methods 0.000 description 1
- 238000005457 optimization Methods 0.000 description 1
- 238000005096 rolling process Methods 0.000 description 1
- 230000007704 transition Effects 0.000 description 1
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C39/00—Aircraft not otherwise provided for
- B64C39/02—Aircraft not otherwise provided for characterised by special use
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C27/00—Rotorcraft; Rotors peculiar thereto
- B64C27/32—Rotors
- B64C27/46—Blades
- B64C27/467—Aerodynamic features
Definitions
- the present invention relates to a rotor blade used in a drone, a helicopter, etc., and a drone and a helicopter having such a rotor blade.
- a drone called a drone with four or more rotor blades is actively used for photographing and observation purposes.
- the drone usually uses a plastic two-bladed fixed-pitch rotor, and the weight of the payload and the flight time are affected by the performance of this rotor blade.
- the rotor blade shape varies, and the chord length of the root is wide, which seems to have applied the classical propeller theory (Non-patent Document 1) from simple taper, and the chord length is sharply reduced from there to the tip. There is a transition to a smaller taper ratio, with rounded tips.
- Non-patent Document 2 developed by GEN CORPORATION as a manned small helicopter, and a simple taper is adopted for the rotor blade.
- Westland Linx in the UK has achieved a world speed record using a rotor blade with an enlarged tip chord length called BERP (Non-patent Document 3, Patent Document 1) and a receding angle. Yes.
- This rotor blade functions as a dog tooth used in an airplane with a discontinuous leading edge with an enlarged chord at the tip, generating a vertical vortex and suppressing peeling of the upper surface, earning a lift coefficient on the reverse side,
- the purpose is to reduce the wave resistance on the forward side by the receding angle of the tip.
- the helicopter has to trade off the maximum speed with the figure of merit (non-patent document 4) obtained by dividing the actual lift by the theoretical maximum lift calculated from the simple motion theory, rotor rotation area and input power.
- a rectangular shape or a shape with a weak taper at the tip is often used.
- the drone since the drone takes a long time to hover, there is a high demand to increase the flight time and the payload by using a rotor blade having a high figure of merit.
- Non-Patent Document 1 There is a method by Adkins et al. (Non-Patent Document 1) as an effective method for designing a rotor blade having a large figure of merit.
- the method of Adkins et al. could not reflect the contraction (reduction in the radius of the cylindrical flow) or the vortex (roll-up) observed in the actual wake of the rotor blade. Therefore, the figure of merit was not the maximum.
- an object of the present invention is to provide a rotor blade capable of increasing the figure of merit, a drone having such a rotor blade, and a helicopter.
- the rotor blade according to the present invention is a rotor blade having a solidity of 10% or less, and the shape of the rotor blade is inversely tapered in a region having a radius of 50% to 90%.
- the maximum chord length value is less than twice the minimum chord length at a radius of 50% to 90%, and the tip gradually narrows The angle at which the base is attached decreases as it goes to the base.
- the drone which concerns on another form of this invention has the rotor blade of the said structure.
- a helicopter according to still another embodiment of the present invention has the rotor blade having the above-described configuration.
- the figure of merit is high according to the present invention, and if used for drones, helicopters, etc., flight time and payload weight can be increased.
- FIG. 1 It is a top view of the rotor blade which concerns on one Embodiment of this invention.
- the trajectory of the blade tip vortex of the rotor blade is shown.
- a) is a trajectory of the blade tip vortex of the rotor blade by the method of Adkins et al.
- b) is a trajectory of the blade tip vortex of the rotor blade by the Harada vortex method of the present inventor.
- FIG. 1 is a top view of a rotor blade according to an embodiment of the present invention. As shown in FIG. 1, the rotor blade 1 is provided on a hub 2 having a shaft hole 3.
- the method of Adkins et al. Can handle only a wake having a constant pitch and a constant radius, as shown in FIG.
- the vortex radius contracts in the same manner as in the experiment as shown in FIG. 2B, and is released from the rotor blade 1 as shown in FIG.
- a discharge vortex model in which the vortex to be concentrated is concentrated at the blade tip and the center of rotation by rolling up behind, a rotor blade 1 having the following form is obtained, and by using such a rotor blade 1 A high figure of merit is obtained.
- the rotor blade 1 has a solidity of 10% or less.
- the shape of the rotor blade 1 is as follows.
- FIG. 4 shows the chord length distribution when the Harada vortex method of the present inventor is optimized using the vortex model of FIG. 2 a), and FIG. 5 shows the mounting angle.
- Table 1 the calculation conditions are shown in Table 1.
- FIG. 4 shows the chord length of the root.
- Some commercially available drones cut the chord length at the base of this rotor blade to about 4 cm.
- FIG. 6 shows the chord length distribution when the Harada vortex method of the present inventor is optimized using the vortex model of FIG. 2 b), and
- FIG. 7 shows the mounting angle. The figure of merit at this time is 78.8%.
- the tangent line AB at the tip of the obtained chord length has a negative slope
- the tangent line CD at a radius of 65% has a positive slope
- the maximum value H is taken at point G.
- This H value is more than twice F.
- the obtained tangent IJ near the tip of the mounting angle has a positive value, takes a minimum value N at a point M near a radius of 60%, takes a maximum value P at a point O near a radius of 30%, and near the base.
- the tangent line KL has a negative value. Even in this shape, the chord length is 7 cm near the base of the wing, which is not realistic.
- the tangent line AB at the tip of the obtained chord length has a negative gradient
- the tangent line CD at a radius of 65% has a positive gradient
- a local minimum at a point E near a radius of 60% Take F and take the maximum value H at point G near the root.
- This H value is more than twice F. Also, the obtained tangent IJ near the tip of the mounting angle has a positive value, takes a minimum value N at a point M near a radius of 60%, takes a maximum value P at a point O near a radius of 30%, and near the base.
- the tangent line KL has a negative value.
- the rotor blade 1 according to this embodiment has a high figure of merit, and if used in a drone, helicopter, etc., the flight time and payload weight can be increased.
- the present invention can be applied to a fixed pitch coaxial double reversing rotor type manned helicopter, a fixed pitch coaxial double reversing rotor type unmanned helicopter, and the like.
- 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.
- 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
- w i ⁇ Z ij ⁇ j (2)
- X ij and Z ij are constants obtained from Biosavart's law
- ⁇ j is the j-th emission vortex.
- FIG. 12 shows the cross section of the blade and the inflow speed.
- FIG. 13 shows the force acting on the blade blade element.
- c i 2 ⁇ i / C L V i (9)
- 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 ⁇ ⁇ V i 2 C D c i db (10)
- 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 cos ⁇ i -dD i sin ⁇ i (11)
- T B ⁇ dT i (14)
- P B ⁇ dP i (15) Given in. Here, B is the number of blades.
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Abstract
[Problem] To be able to provide a high figure of merit and increase flight time and payload weight when used, for example, in a drone or a helicopter. [Solution] A rotor blade 1 has a solidity of 10% or less. The shape of the rotor blade 1 is as follows: (1) the rotor blade has an inversely tapered part in a region from 50% to 90% in radius; (2) the rotor blade has a maximum chord length near the base, and the value thereof is at most twice the minimum value of the chord length within the range from 50% to 90% in radius; (3) the tip end gradually narrows; and (4) the angle of incidence of the base decreases toward the base.
Description
本発明は、ドローンやヘリコプタなどに用いられるロータブレード、そのようなロータブレードを有するドローン及びヘリコプタに関する。
The present invention relates to a rotor blade used in a drone, a helicopter, etc., and a drone and a helicopter having such a rotor blade.
近年、４発以上のロータブレードを持つドローンと呼ばれる飛行体が撮影、観測目的で盛んに利用されている。ドローンは通常プラスチック製の２枚羽の固定ピッチロータを使用しており、このロータブレードの性能によりペイロードの重さ及び飛行時間が左右される。ロータブレード形状は様々であり、単純テーパから古典プロペラ理論（非特許文献１）を応用したと思われる、付け根の翼弦長が広く、そこから先端に向けて急激に翼弦長を減らし、ある程度小さなテーパ比に移行し、先端を丸めた形状のものなどが見られる。
In recent years, a drone called a drone with four or more rotor blades is actively used for photographing and observation purposes. The drone usually uses a plastic two-bladed fixed-pitch rotor, and the weight of the payload and the flight time are affected by the performance of this rotor blade. The rotor blade shape varies, and the chord length of the root is wide, which seems to have applied the classical propeller theory (Non-patent Document 1) from simple taper, and the chord length is sharply reduced from there to the tip. There is a transition to a smaller taper ratio, with rounded tips.
また有人小型ヘリコプタとしてＧＥＮ ＣＯＲＰＯＲＡＴＩＯＮが開発したＧＥＮ Ｈ－４（非特許文献２）同様に固定ピッチブレードを使用しており、ロータブレードに単純テーパを採用している。ヘリコプタの分野では英国のＷｅｓｔｒａｎｄ ＬｉｎｘがＢＥＲＰ（非特許文献３，特許文献１）と呼ばれる先端の翼弦長を拡大し、後退角を持たせたロータブレードを使用し、世界速度記録を達成している。本ロータブレードは先端の翼弦を拡大した不連続な前縁を飛行機で使用されるドッグツースとして機能させ、縦渦を発生させて上面の剥離を抑えることで後退側での揚力係数を稼ぐとともに、先端の後退角によって前進側での造波抵抗を減少させる事を目的としている。
Also, a fixed pitch blade is used as in GEN H-4 (Non-patent Document 2) developed by GEN CORPORATION as a manned small helicopter, and a simple taper is adopted for the rotor blade. In the field of helicopter, Westland Linx in the UK has achieved a world speed record using a rotor blade with an enlarged tip chord length called BERP (Non-patent Document 3, Patent Document 1) and a receding angle. Yes. This rotor blade functions as a dog tooth used in an airplane with a discontinuous leading edge with an enlarged chord at the tip, generating a vertical vortex and suppressing peeling of the upper surface, earning a lift coefficient on the reverse side, The purpose is to reduce the wave resistance on the forward side by the receding angle of the tip.
ヘリコプタは単純運動理論とロータ回転面積と入力パワから求められる理論最大揚力で実際の揚力を除したフィギュアオブメリット（非特許文献４）と最大速度をトレードオフしなければならないため、ロータブレード形状は矩形か先端に弱いテーパがついた形状のものが多く使用されている。これに対してドローンはホバリングする時間が長いためフィギュアオブメリットの高いロータブレードを使用し、飛行時間及びペイロードを増加させようとする要求が高い。
The helicopter has to trade off the maximum speed with the figure of merit (non-patent document 4) obtained by dividing the actual lift by the theoretical maximum lift calculated from the simple motion theory, rotor rotation area and input power. A rectangular shape or a shape with a weak taper at the tip is often used. On the other hand, since the drone takes a long time to hover, there is a high demand to increase the flight time and the payload by using a rotor blade having a high figure of merit.
フィギュアオブメリットの大きいロータブレードを設計する際に有効な手法としてＡｄｋｉｎｓらの方法（非特許文献１）がある。しかしＡｄｋｉｎｓらの方法は実際のロータブレードの後流に見られる縮流（筒状の流れの半径の縮小）や、渦巻き込み（ロールアップ）を反映できなかった。そのためフィギュアオブメリットは最大とはなっていなかった。
There is a method by Adkins et al. (Non-Patent Document 1) as an effective method for designing a rotor blade having a large figure of merit. However, the method of Adkins et al. Could not reflect the contraction (reduction in the radius of the cylindrical flow) or the vortex (roll-up) observed in the actual wake of the rotor blade. Therefore, the figure of merit was not the maximum.
以上のような事情に鑑み、本発明の目的は、フィギュアオブメリットが高くする事ができるロータブレード、そのようなロータブレードを有するドローン及びヘリコプタを提供することにある。
In view of the circumstances as described above, an object of the present invention is to provide a rotor blade capable of increasing the figure of merit, a drone having such a rotor blade, and a helicopter.
上記課題を解決するために、本発明に係るロータブレードは、ソリディティが１０％以下となるロータブレードであり、前記ロータブレードの形状が、半径５０％から９０％の領域で逆テーパとなっている部分を持ち、付け根付近に最大翼弦長を持ち、前記最大翼弦長の値は半径５０％から９０％にある翼弦長の極小値の２倍以下であり、先端は徐々に細くなっており、付け根の取付け角は付け根に行くほど減少している。
In order to solve the above-mentioned problems, the rotor blade according to the present invention is a rotor blade having a solidity of 10% or less, and the shape of the rotor blade is inversely tapered in a region having a radius of 50% to 90%. With a maximum chord length near the root, the maximum chord length value is less than twice the minimum chord length at a radius of 50% to 90%, and the tip gradually narrows The angle at which the base is attached decreases as it goes to the base.
本発明の別形態に係るドローンは、上記構成のロータブレードを有する。
本発明の更に別形態に係るヘリコプタは、上記構成のロータブレードを有する。 The drone which concerns on another form of this invention has the rotor blade of the said structure.
A helicopter according to still another embodiment of the present invention has the rotor blade having the above-described configuration.
本発明の更に別形態に係るヘリコプタは、上記構成のロータブレードを有する。 The drone which concerns on another form of this invention has the rotor blade of the said structure.
A helicopter according to still another embodiment of the present invention has the rotor blade having the above-described configuration.
本発明により、フィギュアオブメリットが高く、ドローン、ヘリコプタなどに使用すれば、飛行時間、ペイロード重量を増す事ができる。
The figure of merit is high according to the present invention, and if used for drones, helicopters, etc., flight time and payload weight can be increased.
以下、図面を参照しながら、本発明の実施形態を説明する。
図１は本発明の一実施形態に係るロータブレードの上面図である。
図１に示すように、ロータブレード１は、シャフト穴３を有するハブ２に設けられている。 Hereinafter, embodiments of the present invention will be described with reference to the drawings.
FIG. 1 is a top view of a rotor blade according to an embodiment of the present invention.
As shown in FIG. 1, therotor blade 1 is provided on a hub 2 having a shaft hole 3.
図１は本発明の一実施形態に係るロータブレードの上面図である。
図１に示すように、ロータブレード１は、シャフト穴３を有するハブ２に設けられている。 Hereinafter, embodiments of the present invention will be described with reference to the drawings.
FIG. 1 is a top view of a rotor blade according to an embodiment of the present invention.
As shown in FIG. 1, the
ここで、Ａｄｋｉｎｓらの方法では図２のａ）に示す、一定ピッチかつ一定半径の後流しか扱えない。これに対して、本発明者である原田の渦法（後述する）では図２のｂ）に示すように渦半径が実験と同様に収縮し、かつ図３に示すようにロータブレード１から放出する渦が後方でロールアップにより翼端と回転中心に集中する放出渦モデルを適用する事で、以下の形態を特徴するロータブレード１を得、そのような形態のロータブレード１を使用する事で高いフィギュアオブメリットを得るものである。
ロータブレード１は、ソリディティが１０％以下である。
ロータブレード１の形状は、以下の通りである。
（１）半径５０％から９０％の領域で逆テーパとなっている部分を持つこと。
（２）付け根付近に最大翼弦長を持ち、その値は半径５０％から９０％にある翼弦長の極小値の２倍以下であること。
（３）先端は徐々に細くなっていること。
（４）付け根の取付け角は付け根に行くほど減少していること。
図４に本発明者である原田の渦法を図２のａ）の渦モデルを用いて最適化した際の翼弦長分布を、また図５に取付角を示す。ここで計算条件を表１に示す。 Here, the method of Adkins et al. Can handle only a wake having a constant pitch and a constant radius, as shown in FIG. On the other hand, in the Harada vortex method (which will be described later), which is the present inventor, the vortex radius contracts in the same manner as in the experiment as shown in FIG. 2B, and is released from therotor blade 1 as shown in FIG. By applying a discharge vortex model in which the vortex to be concentrated is concentrated at the blade tip and the center of rotation by rolling up behind, a rotor blade 1 having the following form is obtained, and by using such a rotor blade 1 A high figure of merit is obtained.
Therotor blade 1 has a solidity of 10% or less.
The shape of therotor blade 1 is as follows.
(1) To have a portion having a reverse taper in a region having a radius of 50% to 90%.
(2) The maximum chord length is near the root, and the value is not more than twice the minimum chord length at a radius of 50% to 90%.
(3) The tip is gradually becoming thinner.
(4) The attachment angle of the base should be reduced toward the base.
FIG. 4 shows the chord length distribution when the Harada vortex method of the present inventor is optimized using the vortex model of FIG. 2 a), and FIG. 5 shows the mounting angle. Here, the calculation conditions are shown in Table 1.
ロータブレード１は、ソリディティが１０％以下である。
ロータブレード１の形状は、以下の通りである。
（１）半径５０％から９０％の領域で逆テーパとなっている部分を持つこと。
（２）付け根付近に最大翼弦長を持ち、その値は半径５０％から９０％にある翼弦長の極小値の２倍以下であること。
（３）先端は徐々に細くなっていること。
（４）付け根の取付け角は付け根に行くほど減少していること。
図４に本発明者である原田の渦法を図２のａ）の渦モデルを用いて最適化した際の翼弦長分布を、また図５に取付角を示す。ここで計算条件を表１に示す。 Here, the method of Adkins et al. Can handle only a wake having a constant pitch and a constant radius, as shown in FIG. On the other hand, in the Harada vortex method (which will be described later), which is the present inventor, the vortex radius contracts in the same manner as in the experiment as shown in FIG. 2B, and is released from the
The
The shape of the
(1) To have a portion having a reverse taper in a region having a radius of 50% to 90%.
(2) The maximum chord length is near the root, and the value is not more than twice the minimum chord length at a radius of 50% to 90%.
(3) The tip is gradually becoming thinner.
(4) The attachment angle of the base should be reduced toward the base.
FIG. 4 shows the chord length distribution when the Harada vortex method of the present inventor is optimized using the vortex model of FIG. 2 a), and FIG. 5 shows the mounting angle. Here, the calculation conditions are shown in Table 1.
図４に示すように付け根の翼弦長が７．５ｃｍに達しているため、実際的ではない。市販のドローンの中にはこのロータブレードの付け根の翼弦長を４ｃｍ程度に切り落として使用している例がある。
図６に本発明者である原田の渦法を図２のｂ）の渦モデルを用いて最適化した際の翼弦長分布を、また図７に取付角を示す。このときのフィギュアオブメリットは７８．８％である。 As shown in FIG. 4, the chord length of the root has reached 7.5 cm, which is not practical. Some commercially available drones cut the chord length at the base of this rotor blade to about 4 cm.
FIG. 6 shows the chord length distribution when the Harada vortex method of the present inventor is optimized using the vortex model of FIG. 2 b), and FIG. 7 shows the mounting angle. The figure of merit at this time is 78.8%.
図６に本発明者である原田の渦法を図２のｂ）の渦モデルを用いて最適化した際の翼弦長分布を、また図７に取付角を示す。このときのフィギュアオブメリットは７８．８％である。 As shown in FIG. 4, the chord length of the root has reached 7.5 cm, which is not practical. Some commercially available drones cut the chord length at the base of this rotor blade to about 4 cm.
FIG. 6 shows the chord length distribution when the Harada vortex method of the present inventor is optimized using the vortex model of FIG. 2 b), and FIG. 7 shows the mounting angle. The figure of merit at this time is 78.8%.
得られた翼弦長の先端の接線ＡＢは負の勾配を持ち、半径６５％付近での接線ＣＤは正の勾配を持ち、半径６０％付近の点Ｅにおいて極小値Ｆをとり、付け根付近の点Ｇで最大値Ｈをとる。
The tangent line AB at the tip of the obtained chord length has a negative slope, the tangent line CD at a radius of 65% has a positive slope, takes a local minimum value F at a point E near a radius of 60%, and is near the root. The maximum value H is taken at point G.
このＨの値はＦの２倍以上である。また得られた取付角の先端付近の接線ＩＪは正の値を持ち、半径６０％付近の点Ｍで極小値Ｎをとり、半径３０％付近の点Ｏで極大値Ｐをとり、付け根付近の接線ＫＬは負の値を持つ。この形状でも翼付け根付近で翼弦長が７ｃｍとなり現実的ではない。
This H value is more than twice F. Also, the obtained tangent IJ near the tip of the mounting angle has a positive value, takes a minimum value N at a point M near a radius of 60%, takes a maximum value P at a point O near a radius of 30%, and near the base. The tangent line KL has a negative value. Even in this shape, the chord length is 7 cm near the base of the wing, which is not realistic.
本発明者である原田の渦法は評価関数を最小化する事で解を得るので翼弦長が４ｃｍを超えるときに値が大きくなるペナルティ関数を評価関数に加えることで翼弦長の制限を課す事ができる。このときの翼弦長を図８に、取付角を図９に示す。このときのフィギュアオブメリットは７８．６％である。
Since Harada's vortex method, which is the present inventor, obtains a solution by minimizing the evaluation function, adding a penalty function that increases when the chord length exceeds 4 cm to limit the chord length. Can be imposed. The chord length at this time is shown in FIG. 8, and the mounting angle is shown in FIG. The figure of merit at this time is 78.6%.
先の結果と同様に、得られた翼弦長の先端の接線ＡＢは負の勾配を持ち、半径６５％付近での接線ＣＤは正の勾配を持ち、半径６０％付近の点Ｅにおいて極小値Ｆをとり、付け根付近の点Ｇで最大値Ｈをとる。
Similar to the previous result, the tangent line AB at the tip of the obtained chord length has a negative gradient, the tangent line CD at a radius of 65% has a positive gradient, and a local minimum at a point E near a radius of 60%. Take F and take the maximum value H at point G near the root.
このＨの値はＦの２倍以上である。また得られた取付角の先端付近の接線ＩＪは正の値を持ち、半径６０％付近の点Ｍで極小値Ｎをとり、半径３０％付近の点Ｏで極大値Ｐをとり、付け根付近の接線ＫＬは負の値を持つ。
This H value is more than twice F. Also, the obtained tangent IJ near the tip of the mounting angle has a positive value, takes a minimum value N at a point M near a radius of 60%, takes a maximum value P at a point O near a radius of 30%, and near the base. The tangent line KL has a negative value.
従って、この実施形態に係るロータブレード１は、フィギュアオブメリットが高く、ドローン、ヘリコプタなどに使用すれば、飛行時間、ペイロード重量を増す事ができる。
本発明は、固定ピッチ同軸２重反転ロータ式有人ヘリコプタや固定ピッチ同軸２重反転ロータ式無人ヘリコプタなどに適用する事ができる。 Therefore, therotor blade 1 according to this embodiment has a high figure of merit, and if used in a drone, helicopter, etc., the flight time and payload weight can be increased.
The present invention can be applied to a fixed pitch coaxial double reversing rotor type manned helicopter, a fixed pitch coaxial double reversing rotor type unmanned helicopter, and the like.
本発明は、固定ピッチ同軸２重反転ロータ式有人ヘリコプタや固定ピッチ同軸２重反転ロータ式無人ヘリコプタなどに適用する事ができる。 Therefore, the
The present invention can be applied to a fixed pitch coaxial double reversing rotor type manned helicopter, a fixed pitch coaxial double reversing rotor type unmanned helicopter, and the like.
次に、上述した原田の渦法を説明する。
図１０にプロペラの座標系と放出渦を示す。プロペラは回転しながらｘ軸方向に移動し、移動した軌跡に放出渦が残されると考える。
図１１に一枚目のブレードの拡大図を示す。回転軸からｒｉ離れたブレード上にｉ番目の代表点があり、ｊ番目の放出渦は図中白丸で示された様に短い線分に離散化する。ｊ番目の単位強度の放出渦がｉ番目の代表点に引き起こすｘ方向の誘導速度をＸｉｊ、ｚ方向の誘導速度をＺｉｊとすると、ｉ番目の代表点に引き起こされるｘ方向の誘導速度ｕｉ、ｚ方向の誘導速度ｗｉはそれぞれ
ｕｉ＝ΣＸｉｊгｊ （１）
ｗｉ＝ΣＺｉｊгｊ （２）
で与えられる。ここで、Ｘｉｊ、Ｚｉｊはビオサバールの法則から得られる定数であり、гｊはｊ番目の放出渦である。 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 (1)
w i = ΣZ ij г j (2)
Given in. Here, X ij and Z ij are constants obtained from Biosavart's law, and г j is the j-th emission vortex.
図１０にプロペラの座標系と放出渦を示す。プロペラは回転しながらｘ軸方向に移動し、移動した軌跡に放出渦が残されると考える。
図１１に一枚目のブレードの拡大図を示す。回転軸からｒｉ離れたブレード上にｉ番目の代表点があり、ｊ番目の放出渦は図中白丸で示された様に短い線分に離散化する。ｊ番目の単位強度の放出渦がｉ番目の代表点に引き起こすｘ方向の誘導速度をＸｉｊ、ｚ方向の誘導速度をＺｉｊとすると、ｉ番目の代表点に引き起こされるｘ方向の誘導速度ｕｉ、ｚ方向の誘導速度ｗｉはそれぞれ
ｕｉ＝ΣＸｉｊгｊ （１）
ｗｉ＝ΣＺｉｊгｊ （２）
で与えられる。ここで、Ｘｉｊ、Ｚｉｊはビオサバールの法則から得られる定数であり、гｊはｊ番目の放出渦である。 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 (1)
w i = ΣZ ij г j (2)
Given in. Here, X ij and Z ij are constants obtained from Biosavart's law, and г j is the j-th emission vortex.
図１２にブレードの断面と流入速度を示す。ブレードに流入する空気の接線方向成分をＵＴとするとＵＴは
ＵＴ＝ｒｉΩ－ｗｉ （３）
で与えられる。ここでΩはプロペラの回転角速度である。またブレードに流入する空気の軸方向成分をＵＰとするとＵＰは
ＵＰ＝Ｕ－ｕｉ （４）
で与えられる。ここでＵはプロペラの前進速度である。流入角φｉ及び流入速度Ｖｉはそれぞれ次式で与えられる。
φｉ＝ｔａｎ－１（ＵＰ／ＵＴ） （５）
Ｖｉ＝√（ＵＰ ２＋ＵＴ ２） （６） 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 (3)
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 ( 4)
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) (5)
V i = √ (U P 2 + U T 2) (6)
ＵＴ＝ｒｉΩ－ｗｉ （３）
で与えられる。ここでΩはプロペラの回転角速度である。またブレードに流入する空気の軸方向成分をＵＰとするとＵＰは
ＵＰ＝Ｕ－ｕｉ （４）
で与えられる。ここでＵはプロペラの前進速度である。流入角φｉ及び流入速度Ｖｉはそれぞれ次式で与えられる。
φｉ＝ｔａｎ－１（ＵＰ／ＵＴ） （５）
Ｖｉ＝√（ＵＰ ２＋ＵＴ ２） （６） 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 (3)
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 ( 4)
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) (5)
V i = √ (U P 2 + U T 2) (6)
図１３にブレード翼素に働く力を示す。ｉ番目の翼素に働く局所揚力ｄＬｉはクッタ=ジューコフスキーの定理により
ｄＬｉ＝ρＶｉгｉｄｂ （７）
で与えられる。ここでρは空気密度、ｄｂは翼素の幅である。または、揚力係数ＣＬを用いて
ｄＬｉ＝１／２・ρＶｉ ２ＣＬｃｉｄｂ （８）
で表される。ここで、ｃｉはｉ番目の翼素の翼弦長である。（７）、（８）式よりｃｉは次式で与えられる。
ｃｉ＝２гｉ／ＣＬＶｉ （９）
ｉ番目の翼素に働く局所抵抗ｄＤｉは抵抗係数ＣＤを用いて次式で与えられる。
ｄＤｉ＝１／２・ρＶｉ ２ＣＤｃｉｄｂ （１０）
ＣＤはレイノルズ数とＣＬの関数であるが、ＣＬに定数を用い、ＣＤがレイノルズ数に対しては鈍感であるとして定数としてよい。
局所揚力ｄＬｉと局所抵抗ｄＤｉの合力の軸方向分力は局所推力ｄＴｉとなり、接線方向分力をｄＮｉとするとそれぞれ、
ｄＴｉ＝ｄＬｉｃｏｓφｉ－ｄＤｉｓｉｎφｉ （１１）
ｄＮｉ＝ｄＬｉｓｉｎφｉ＋ｄＤｉｃｏｓφｉ （１２）
となる。局所吸収パワｄＰｉはｄＮｉｒｉΩであるから次式で与えられる。
ｄＰｉ＝（ｄＬｉｓｉｎφｉ＋ｄＤｉｃｏｓφｉ）ｒｉΩ （１３）
結局、推力と吸収パワは（１１）式と（１３）式よりそれぞれ
Ｔ＝ＢΣｄＴｉ （１４）
Ｐ＝ＢΣｄＰｉ （１５）
で与えられる。ここでＢはブレード枚数である。 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 (7)
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 (8)
It is represented by Here, c i is the chord length of the i th blade element. (7), c i from equation (8) is given by the following equation.
c i = 2г i / C L V i (9)
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 (10)
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 (11)
dN i = dL i sin φ i + dD i cos φ i (12)
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 (13)
Eventually, the thrust and absorption power are calculated from the equations (11) and (13) respectively. T = BΣdT i (14)
P = BΣdP i (15)
Given in. Here, B is the number of blades.
ｄＬｉ＝ρＶｉгｉｄｂ （７）
で与えられる。ここでρは空気密度、ｄｂは翼素の幅である。または、揚力係数ＣＬを用いて
ｄＬｉ＝１／２・ρＶｉ ２ＣＬｃｉｄｂ （８）
で表される。ここで、ｃｉはｉ番目の翼素の翼弦長である。（７）、（８）式よりｃｉは次式で与えられる。
ｃｉ＝２гｉ／ＣＬＶｉ （９）
ｉ番目の翼素に働く局所抵抗ｄＤｉは抵抗係数ＣＤを用いて次式で与えられる。
ｄＤｉ＝１／２・ρＶｉ ２ＣＤｃｉｄｂ （１０）
ＣＤはレイノルズ数とＣＬの関数であるが、ＣＬに定数を用い、ＣＤがレイノルズ数に対しては鈍感であるとして定数としてよい。
局所揚力ｄＬｉと局所抵抗ｄＤｉの合力の軸方向分力は局所推力ｄＴｉとなり、接線方向分力をｄＮｉとするとそれぞれ、
ｄＴｉ＝ｄＬｉｃｏｓφｉ－ｄＤｉｓｉｎφｉ （１１）
ｄＮｉ＝ｄＬｉｓｉｎφｉ＋ｄＤｉｃｏｓφｉ （１２）
となる。局所吸収パワｄＰｉはｄＮｉｒｉΩであるから次式で与えられる。
ｄＰｉ＝（ｄＬｉｓｉｎφｉ＋ｄＤｉｃｏｓφｉ）ｒｉΩ （１３）
結局、推力と吸収パワは（１１）式と（１３）式よりそれぞれ
Ｔ＝ＢΣｄＴｉ （１４）
Ｐ＝ＢΣｄＰｉ （１５）
で与えられる。ここでＢはブレード枚数である。 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 (7)
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 (8)
It is represented by Here, c i is the chord length of the i th blade element. (7), c i from equation (8) is given by the following equation.
c i = 2г i / C L V i (9)
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 (10)
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 (11)
dN i = dL i sin φ i + dD i cos φ i (12)
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 (13)
Eventually, the thrust and absorption power are calculated from the equations (11) and (13) respectively. T = BΣdT i (14)
P = BΣdP i (15)
Given in. Here, B is the number of blades.
プロペラの最適設計問題はＣＬ、ＣＤ、Ω、設計パワＰ０を設定し、гｉを未知数として、
［条件］：Ｐ＝Ｐ０
のもとで
［目的関数］：－Ｔ
を最小化する
最適化問題に帰着する。 The optimal design problem for the propeller is C L , C D , Ω, design power P 0 , and гi as an unknown,
[Condition]: P = P 0
[Objective function]: -T
This results in an optimization problem that minimizes.
［条件］：Ｐ＝Ｐ０
のもとで
［目的関数］：－Ｔ
を最小化する
最適化問題に帰着する。 The optimal design problem for the propeller is C L , C D , Ω, design power P 0 , and гi as an unknown,
[Condition]: P = P 0
[Objective function]: -T
This results in an optimization problem that minimizes.
１ ロータブレード
２ ハブ
３ シャフト穴 1Rotor blade 2 Hub 3 Shaft hole
２ ハブ
３ シャフト穴 1
Claims (3)
- ソリディティが１０％以下となるロータブレードであり、
前記ロータブレードの形状が、
半径５０％から９０％の領域で逆テーパとなっている部分を持ち、
付け根付近に最大翼弦長を持ち、前記最大翼弦長の値は半径５０％から９０％にある翼弦長の極小値の２倍以下であり、
先端は徐々に細くなっており、
付け根の取付け角は付け根に行くほど減少している
ロータブレード。 A rotor blade with a solidity of 10% or less,
The shape of the rotor blade is
It has a part that is inversely tapered in the radius 50% to 90% region,
It has a maximum chord length near the root, and the value of the maximum chord length is less than twice the minimum value of the chord length at a radius of 50% to 90%,
The tip is gradually getting thinner
The mounting angle of the base decreases as it goes to the base. - 請求項１に記載のロータブレードを有するドローン。 A drone having the rotor blade according to claim 1.
- 請求項１に記載のロータブレードを有するヘリコプタ。 A helicopter having the rotor blade according to claim 1.
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|---|---|---|---|---|
| US20180127086A1 (en) * | 2016-11-10 | 2018-05-10 | Coretronic Intelligent Robotics Corporation | Aerial Vehicle and Propeller Thereof |
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| JP6856930B2 (en) * | 2017-02-15 | 2021-04-14 | 国立研究開発法人宇宙航空研究開発機構 | Rotor, drone and helicopter |
| US20230257113A1 (en) | 2020-06-30 | 2023-08-17 | Sony Group Corporation | Propeller, flying object, and method for manufacturing propeller |
Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB752080A (en) * | 1952-07-18 | 1956-07-04 | Aristides Warto | Improvements in or relating to rotor blades for helicopters |
| US5174721A (en) * | 1990-10-13 | 1992-12-29 | Westland Helicopters Limited | Helicopter rotor blades |
| JP2009173152A (en) * | 2008-01-24 | 2009-08-06 | Mitsubishi Heavy Ind Ltd | Helicopter, its rotor, and its control method |
-
2017
- 2017-01-13 WO PCT/JP2017/001030 patent/WO2017145563A1/en active Application Filing
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Patent Citations (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB752080A (en) * | 1952-07-18 | 1956-07-04 | Aristides Warto | Improvements in or relating to rotor blades for helicopters |
| US5174721A (en) * | 1990-10-13 | 1992-12-29 | Westland Helicopters Limited | Helicopter rotor blades |
| JP2009173152A (en) * | 2008-01-24 | 2009-08-06 | Mitsubishi Heavy Ind Ltd | Helicopter, its rotor, and its control method |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US20180127086A1 (en) * | 2016-11-10 | 2018-05-10 | Coretronic Intelligent Robotics Corporation | Aerial Vehicle and Propeller Thereof |
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