WO2009096010A1 - Propeller aircraft, propeller apparatus, and posture controller - Google Patents

Propeller aircraft, propeller apparatus, and posture controller Download PDF

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Publication number
WO2009096010A1
WO2009096010A1 PCT/JP2008/051415 JP2008051415W WO2009096010A1 WO 2009096010 A1 WO2009096010 A1 WO 2009096010A1 JP 2008051415 W JP2008051415 W JP 2008051415W WO 2009096010 A1 WO2009096010 A1 WO 2009096010A1
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WIPO (PCT)
Prior art keywords
propeller
stabilizer
wing
cylindrical
airframe
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PCT/JP2008/051415
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French (fr)
Japanese (ja)
Inventor
Hiroshi Kawaguchi
Original Assignee
Kawaguchi, Yasuko
Kawaguchi, Syuichi
Kawaguchi, Megumi
Kawaguchi, Sachiko
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
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Application filed by Kawaguchi, Yasuko, Kawaguchi, Syuichi, Kawaguchi, Megumi, Kawaguchi, Sachiko filed Critical Kawaguchi, Yasuko
Priority to PCT/JP2008/051415 priority Critical patent/WO2009096010A1/en
Priority to PCT/JP2008/059529 priority patent/WO2009096048A1/en
Priority to JP2009551392A priority patent/JP5184555B2/en
Priority to PCT/JP2008/065873 priority patent/WO2009096058A1/en
Publication of WO2009096010A1 publication Critical patent/WO2009096010A1/en

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    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C11/00—Propellers, e.g. of ducted type; Features common to propellers and rotors for rotorcraft

Definitions

  • the present invention relates to a propeller aircraft capable of stable vertical take-off and landing and hovering and application of the technology.
  • Patent Documents 1 to 8 have already been published as prior art documents related to the background art of the present invention.
  • An object of the present invention is, firstly, to provide a propeller aircraft capable of stable vertical takeoff and landing or hovering, and secondly, to provide a propeller device capable of reducing a burden for ensuring static stability. Thirdly, it is to provide an attitude control device having excellent stability by applying them.
  • the aircraft includes a plurality of vertical main wings assembled radially, and when the diameter of the propeller is r 0 , the propeller has a moment received from ambient air.
  • the shape of the vertical main wing is formed so that a moment of 2 ⁇ r 0 is applied to the plurality of vertical main wings by the propeller wake.
  • the total moment received by the propeller from the surrounding air and the total moment caused by the propeller wake applied to the aircraft can be made equal, thereby preventing the counter-rotation due to the reaction of the propeller rotation in the aircraft.
  • the airframe includes a plurality of vertical main wings that are radially assembled, and each of the vertical main wings is formed to have the same size as the spread of the propeller wake.
  • a propeller device includes an airframe having a plurality of vertically partitioned vertical partition plates, and a propeller disposed on the upper side of the plurality of vertical partition plates.
  • the shape of the vertical partition plate is formed such that when the diameter is r 0 , a moment 2 ⁇ r 0 times the moment the propeller receives from the surrounding air is applied to the plurality of vertical partition plates by the wake of the propeller. Is.
  • the propeller since the shape of the vertical partition plate is formed such that a moment 2 ⁇ r 0 times the moment the propeller receives from the surrounding air is applied to the plurality of vertical partition plates by the propeller wake, the propeller is The total moment received from the air and the total moment due to the propeller wake applied to the aircraft can be made equal, thereby counteracting the counter torque caused by the propeller rotation reaction in the aircraft. Therefore, the burden for ensuring static stability can be reduced.
  • the vertical partition plates are formed to have the same width as the spread of the propeller wake, and the height ⁇ l of the vertical main wings is defined such that the diameter of the propeller is r 0.
  • ⁇ l 2 ⁇ r 0 / m is set.
  • the airframe further includes a cylindrical body that is covered around the plurality of vertical partition plates.
  • the wake of the propeller can be concentrated and ejected behind the propeller.
  • the airframe is configured to be divided into n (n: an integer of 2 or more) layers, and each of the layers includes a plurality of vertical partition plates assembled radially, When i-th of the height of the layer and the number of diameter unit of a vertical partition plate respectively and .DELTA.l i and m i, height .DELTA.l 1 of each layer, ⁇ l 2, ..., the number m 1 of .DELTA.l n and vertical partition plate , m 2 ,, ..., m n are those set so as to satisfy the ⁇ l 1 m 1 + ⁇ l 2 m 2 + ... + ⁇ l n m n ⁇ 2 ⁇ r 0.
  • the height of the airframe can be easily adjusted simply by adding layers.
  • the height of each layer ⁇ l 1, ⁇ l 2 ,, ..., number m 1, m 2 ,, ... diameter units .DELTA.l n and the vertical partition plate, m n is, ⁇ l 1 m 1 + ⁇ l 2 m 2 + ... Since it is set so as to satisfy + ⁇ l n m n ⁇ 2 ⁇ r 0 , it is possible to cancel the counter torque due to the propeller rotation reaction in the airframe.
  • each of the layers further includes a cylindrical body covered around the plurality of vertical partition plates.
  • the propeller wind can be effectively injected to the rear of the propeller.
  • the attitude control device includes radial stabilizers or cylindrical stabilizers.
  • the central stabilizer direction acts as a resistance against radial shaking of the radial stabilizer wing or the cylindrical stabilizer wing, and the stability of the attitude control device (and thus the airplane equipped with this attitude control device) can be dramatically improved.
  • An attitude control device includes a cylindrical stabilizer, and one or more radial stabilizers arranged coaxially along a central axis of the cylindrical stabilizer. Is.
  • the cylindrical stabilizer blade can prevent the wind flow to the radial stabilizer blade from spreading to the outside of the cylindrical stabilizer blade and can make the wind flow uniform. Stability can be improved. Therefore, the stability at the time of flight of the airplane provided with this attitude control device can be improved.
  • the attitude control device further includes the cylindrical stabilizer and one or more in-cylinder stabilizers coaxially arranged therein.
  • An attitude control device further includes a wind flow generating device that is disposed on a central axis of the cylindrical stabilizer and / or the radial stabilizer and generates a wind flow in the direction of the central axis. It is.
  • the attitude control device can be used as a propulsion device, and stable flight can be performed.
  • the radial stabilizer wing and / or the cylindrical stabilizer wing has a distance n GC between the center of gravity of the attitude control device and the wind pressure center point of each stabilizer wing.
  • the relationship between the center of gravity and the distance n GW between the external wind pressure center point W is arranged as represented by Expression 19- (5).
  • the attitude control device can be stably hovered.
  • the attitude control device is the central axis of the cylindrical stabilizer blade at a position of a distance of 1/8 or more of the length of the cylindrical stabilizer blade downward from the upper end of the cylindrical stabilizer blade.
  • a wind flow generator for generating a wind flow in the direction is arranged.
  • the position of the radial stabilizer can be arbitrarily selected without changing the position of the center of gravity where the attitude control device stably hovers, and the degree of freedom in designing the attitude control device can be improved.
  • the wind flow generator includes a propeller that generates a wind flow and a propeller drive unit.
  • the attitude control device is such that an auxiliary stabilizer blade is disposed under the cylindrical stabilizer blade.
  • the overall wind pressure center point of the entire attitude control device is lowered, so that the position of the center of gravity for stably hovering the attitude control device is also lowered, and the restoration effect by the center of gravity is increased. However, it can be further stabilized.
  • the effective angle of the propeller is ⁇ T
  • the diameter of the stable blade of the i-th radial stabilizer blade is When the number of units is m i and the value obtained by dividing the length of the i-th radial stabilizer blade in the central axis direction by the diameter of the propeller is n i , Equation 14- (1) is established.
  • the counter torque due to the propeller rotation can be offset, and the attitude control device can be prevented from counter rotating due to the propeller rotation.
  • a posture control device is a posture control device that is a combination of two of the eleventh to sixteenth posture control devices, each of which has its intake side open end on the upper side. And the exhaust-side opening end facing downward, and the intake-side opening ends are inclined in directions opposite to each other (including an inclination angle of 0 °).
  • FIG. 4 is a sectional view taken along line VI-VI in FIGS. 1 and 3. It is a figure explaining how to obtain
  • FIG. It is an example figure of a propeller.
  • FIG. It is an example figure of the propeller apparatus 50B which concerns on Embodiment 2.
  • FIG. It is the side view and top view of a fuselage of a triangular wing. It is a top view of six radial stabilizers.
  • FIG. 3 is a schematic side view of the airframe when the inclination of the propeller rotation shaft is ⁇ and the angle of blur at the tip of the propeller rotation shaft is ⁇ .
  • FIG. FIG. 6 is a diagram showing a state in which a propeller wake pk acts on a stable blade 61. It is the perspective view and side view of an example of the airframe provided with the cylindrical stabilizer blade and the radial stabilizer blade. It is a perspective view of another example of the airframe provided with the cylindrical stabilizer blade and the radial stabilizer blade. It is the perspective view and top view of further another example of the airframe provided with the cylindrical stabilizer blade and the radial stabilizer blade. It is an example figure of the airframe which combined two airframes of FIG. It is a figure explaining the conditions for canceling out and stabilizing the influence by the shake of a propeller rotating shaft. It is a perspective view of the airframe 80a when the auxiliary stabilizing wing 83b is attached below the airframe 80. FIG.
  • the propeller aircraft 40A of this embodiment is arranged at the upper end of an airframe 41 in which two vertical main wings (hereinafter referred to as main wings) 41a are radially and parallel to each other. And a main propeller 43.
  • Each main wing 41a is formed in a half trapezoidal plate shape of the same shape and size, and the entire body 41 is formed in a trapezoidal plate shape (ie, a delta wing).
  • the inclination angle ⁇ of the outer side of each main wing 41a is an angle that matches the spread of the wind from the propeller 43 during hovering (hereinafter referred to as the propeller wake).
  • point P in FIG. 1 is a vertex of a triangle formed by virtually complementing the trapezoid of the airframe 41, and that all the winds that follow the propeller are generated from this point P. .
  • the length between the point P and the propeller 43 is the length between the point P and the propeller 43
  • the symbol R is the length between the point P and the lower side of the aircraft 41
  • the symbol l is the length between the propeller 43 and the propeller 43.
  • the length between the lower side of the fuselage 41 and the sign ⁇ l is the length of the fuselage 41 (more specifically, the length of the propeller lower part (main wing part) of the fuselage 41)
  • the sign r 0 is The propeller diameter
  • the symbol r l is the width of the airframe 41 at a distance l away from the propeller 43 (more specifically, the width along the main wing surface between two adjacent main wings)
  • the symbols ⁇ 0 and ⁇ l are the air flow density just below the propeller 43 and the air flow density at a point away from the propeller 43 by the distance l, respectively.
  • the height ⁇ l of the lower propeller portion (main wing portion) of the fuselage 41 is sufficiently shorter than the final reach distance of the propeller wake, so the propeller wake flowing on the surface of the fuselage 41 is assumed.
  • the wind speed of can be regarded as almost constant.
  • Equation (1) Since the airflow density ⁇ due to the propeller wake at the point R away from the point P is inversely proportional to R 2 , Equation (1) is obtained.
  • the airframe 41 has a triangular stable wing 41c on the upper side of the propeller 43, and the combined shape of the main wing 41a and the stable wing 41c is a triangle. It is formed to become.
  • a force b (b: vector) is given to the air with respect to a force a (a: vector) received by the propeller 43 by the air hitting the propeller 43, and a force F having a magnitude of the horizontal component asin ⁇ of the force a is This is related to the rotational moment of the airframe 41 around the propeller rotation axis.
  • all the magnitudes of the force represent the total amount of the force.
  • the air receiving the force b becomes wind and finally hits the main wing 41a (propeller lower portion) of the fuselage 41, and the hit angle ⁇ can be regarded as substantially equal to the angle ⁇ of the force a received by the propeller 43.
  • Equation (15) the moment M 0 around the propeller rotation axis received by the propeller 43 is expressed by Equation (15).
  • the total moment [M 0 ] of the propeller 43 is considered.
  • the wind from the propeller 43 is intermittently descended under the airframe 41 when the propeller 43 becomes substantially parallel to the airframe 41.
  • the propeller 43 is not parallel to the main wing 41a, the moment Mo 0 of each moment of the propeller 43 is not transmitted to the main wing 41a.
  • the rotation speed of the propeller 43 is extremely high, it is considered that the time from one rotation to the next parallel time is very short, and the moment obtained by adding all the moments for one rotation may be considered as the total moment.
  • Equation (18) the total moment [M 0 ] of the propeller 43 is expressed by Equation (18).
  • Formula (20) becomes Formula (45) when there are m main wings 41a.
  • the number “2” on the rightmost side of equation (18) is doubled for the reason that the propeller 43 has two blades.
  • the propeller 43 has two blades. Basically, when the propeller 43 is parallel to the airframe 41, the airflow and moment of the entire airframe 41 are calculated on the assumption that the two blades are parallel to the airframe 41 at the same time. This is because the total moment [M 0 ] must be the total moment of the two blades.
  • the equation (45) does not change. This is because the air flow density on the airframe 41 increases as the number of propellers 43 increases. Therefore, even in the case of a three-blade or four-blade propeller 43 as shown in FIGS.
  • the airframe 41 is made, and the center of gravity is arranged at the equilibrium point between the wind pressure center point C calculated by the above equation (64) and the center point by the external wind pressure (hereinafter referred to as the external wind pressure center point). Then, when performing vertical take-off and landing and hovering, it was proved that neither anti-rotation nor left-right shaking occurred at all. 3 is provided for adjusting the position of the external wind pressure center point.
  • a moment [M 0 ] 2 ⁇ r 0 times the moment M 0 received by the propeller 43 from the surrounding air is applied to the plurality of main wings (vertical main wings) 41a by the wake of the propeller. Since the shape of the vertical main wing 41a is formed so as to be applied, the total moment [M 0 ] received by the propeller from the surrounding air can be made equal to the total moment [M l ] due to the propeller wake applied to the airframe 41. Anti-rotation due to the reaction of propeller rotation in the airframe 41 can be stopped.
  • Runode, the height .DELTA.l vertical wing 41a, ⁇ l 2 ⁇ r 0 / m easily by simply set to satisfy the, can stop the counter-rotation by the reaction of the propeller rotation in the aircraft 41.
  • the steering wing, the control unit, the drive unit, the power source, and the like are not particularly described, but are naturally provided in the airframe.
  • a cylindrical tubular body 47 i having the same height as the partition plates 41 a i is provided.
  • the tubular body 47 i is, for example, fixed to the side end surface of the partition plate 41a i.
  • the layers H i are connected and fixed to each other adjacent in the vertical direction, for example, by a connecting member (not shown) via the peripheral surface of each cylindrical body 47 i .
  • each layer H i when the layers H i are arranged vertically one row concentrically propeller shaft 43a, the lower surface 47a i of each layer H i coincides with the boundary line Q spread of the propeller slipstream It is formed like this.
  • Each layer H i is also spaced apart from one another, be arranged without an interval from each other, may either. In the case where the lower surface 47a i of each layer H i protrudes outwardly from the boundary line Q spread of the propeller slipstream is necessary fine adjustment to the slightly smaller height .DELTA.l i of each layer H i.
  • equation (60) becomes 4 ⁇ l 1 + 3 ⁇ l 2 + 6 ⁇ l 3 ⁇ 2 ⁇ r 0 , and this relationship
  • the heights ⁇ l 1 , ⁇ l 2 , ⁇ l 3 of each layer H 1 , H 2 , H 3 may be set so as to satisfy the above.
  • the height of the fuselage 41B (the height of the partition plate portion) can be easily adjusted simply by adding layers. it can. At that time, the height of each layer ⁇ l 1, ⁇ l 2, ..., the number m 1, m 2 ,, ... diameter units .DELTA.l n and the vertical partition plate, since the m n is set to satisfy equation (60) As in the case of the first embodiment, the counter torque caused by the propeller rotation reaction in the airframe 41B can be canceled.
  • the respective layers H i so comprises a tubular body 47 i which Kabusare around the plurality of vertical partition plates 41a i, can cause the propeller wind effectively injected behind a propeller.
  • the fuselage 63 has, for example, a propeller 60, a trapezoidal shape in a side view arranged on the lower side of the propeller 60, and a radial (for example, cross-shaped) lower stabilizer blade 61, and a triangular shape in a side view arranged on the upper side of the propeller 60.
  • a radial (for example, cross-shaped) upper stabilizer blade 62 and a propeller drive unit (not shown) disposed on the lower radial stabilizer blade 61 are provided.
  • G center of gravity of the aircraft 60
  • C wind center point W by propeller backwash streams: external air by wind pressure center point r 0: Propeller diameter n a r 0: distance n b r between the upper side of the propeller 60 and the lower stable wing 61 0 : distance between the propeller 60 and the bottom of the lower stabilizing blade 61
  • n c r 0 distance from the propeller 60 at the wind pressure center point C
  • n W r 0 distance from the propeller 60 at the wind pressure center point W
  • n c 1.304 from Equation 1- (18) and Equation 1- (1).
  • Equation 1- (2) is obtained.
  • the area of one sheet of one sheet of the area S C and 1 sheet of the area and the lower stable wing 61 of the aircraft 63 total projected area (i.e. the upper stable wing 62 of the lower stabilizing wing 61 of the aircraft 63
  • Equation 1- (9) The meaning of Equation 1- (9) is that the force F C caused by the propeller wake acting on the area S C of the stabilizing blade 61 where the propeller wake flows is the ⁇ of the external wind pressure F W ′ applied to the same area. It is to say that it is double. Therefore, the aircraft 63 total projected area S W is, when a W times the area S C of the stable wing 61 the propeller slipstream is flowing, wherein 1- (10) as a general formula and the formula 1- (11) Can be represented.
  • the F C weighs mg aircraft 63 (m: mass of the aircraft 63, g: gravitational acceleration) should be a force proportional to. Therefore, when the proportionality coefficient is K, F C and F W are considered to be expressed by Expression 1- (12) and Expression 1- (13), respectively.
  • Formula 1- (12), Formula 1- (13), and Formula 1- (14) originates from the fact that the height of the lower stabilizer blade 61 of the fuselage 63 is ⁇ times the diameter r 0 of the propeller 60. You can easily guess what you are doing. Therefore, the general formulas of these formulas (that is, formulas in the case where the height of the lower stabilizer blade 61 of the fuselage 63 is n times the diameter r 0 of the propeller 60) are formulas 1- (15), And Formula 1- (17).
  • the stabilizer blade 61-1 is stabilized at the wind pressures F C-2 to F C-6 of the stabilizer blades 61-2 to 61-6.
  • a component perpendicular to the wing 61-1 is applied.
  • the total force [F C ] of the wind pressure applied to the stable blade 61-1 is expressed by Equation 2- (1).
  • Equation 2- (2) becomes Equation 2- (3).
  • N is expressed by Equation 2- (4), where m is the diameter unit number of the stabilizing blade of the lower stabilizing blade (radial stabilizing blade) 61.
  • N 2.
  • Equation 2- (6) the number m of the stable blade diameter unit of the lower stabilizer blade 61 is expressed by Equation 2- (5), Equation 2- (6) always holds.
  • the reason that the airframe 63 is unstable is the wake of the propeller. 2.
  • the force of F C can be assumed to be a force (pseudo lift) similar to lift because the wake behind the propeller flows almost parallel to the lower stabilizer blade 61.
  • the fact that the pseudo-lift force is applied to the lower stabilizing blade 61 in the parallel flow means that the wake behind the propeller considered to be a parallel flow is actually not parallel to the lower stabilizing blade 61 but at an angle. What will happen 4.
  • the reason why the propeller wake has an angle with respect to the lower stabilizer blade 61 is that the rotating shaft of the propeller 60 was originally inclined at a certain angle with respect to the lower stabilizer blade 61 or the tip of the rotating shaft due to the rotation of the propeller 60. Be considered to be blurring.
  • F C is assumed to be a pseudo lift, and the definition of F C is corrected.
  • Formula 4- (1) is known as one of the general formulas of lift L.
  • Equations 4- (2) and 4- (3) are almost correct. Furthermore, from these experiments, it is proved that the pseudo lift coefficient k is k ⁇ 1.
  • FIG. 8 is a schematic side view of the airframe 63 when the inclination of the propeller rotation shaft, which is an unstable element during hovering of the airframe 63, is ⁇ , and the angle of blurring of the tip of the propeller rotation shaft during propeller rotation is ⁇ . is there.
  • Equation 5 When aircraft 63 in windless is hovering, the moment balance equation about the gravity G of the pseudo-lift F C and the propeller thrust F P and related body 63 becomes Equation 5 (9).
  • Equation 5- (9) becomes Equation 5- (1).
  • the center of gravity G of the body 63 and the external wind pressure center point W are matched to determine the position of the center of gravity G that is stable, and from the body 63 at that time, n GC , N C , N, n, cos ( ⁇ + ⁇ ) are obtained, and those values are substituted into the equation 5- (1) to obtain the pseudo lift coefficient k.
  • the numerator on the right side of the equation 5- (1) is n C.
  • n C in Expression 5- (1) can be replaced with n X.
  • the expression 5- (1) (that is, the general expression) becomes the expression 5- (2).
  • Equation 5- (3) Equation 5- (4)
  • Equation 5- (5) Is established.
  • Equation 5- (6) Equation 5- (7)
  • Equation 5- (8) Equation 5- (8)
  • the force F C is a resistance force
  • the position of the center of gravity G of the airframe 63 and the wind pressure center point C are reversed in the experiment of S5 described above (that is, the center of gravity G is disposed above the wind pressure center point C).
  • the orientation of the F C should be reversed accordingly.
  • an equation corresponding to Equation 5- (2) is obtained from the moment balance equation in that case, and the same hovering experiment is performed under the same condition using the equation, the same result (ie, the airframe 63 will be stabilized). Result) should be obtained.
  • Equation 6- (1) is obtained as an equation corresponding to 5- (2).
  • the force F C is a resistance force that is generated regardless of the direction in which the airframe 63 moves, regardless of the direction. This is considered to be the reason why the aircraft that has been stabilized in the experiment is becoming unstable again. Therefore, the n GW corresponding to the n GC of this experiment is obtained from the equation 5- (5), the point W is arranged at the position of the distance n GW r 0 determined by the n GW , and the hovering experiment is performed again. 63 stably hovered.
  • the force of F C is a resistance force when the airframe 63 tries to move. Therefore, it can be seen that no force is generated when the airframe 63 is not moving at all. If F C becomes the force is resistant, F C becomes a resistance to external air, so that the stability of the machine body 63 is increased with respect to the external air during hovering. When the airframe 63 is pushed by the external wind and starts to move, a resistance force F C is generated, and the acceleration of the movement is reduced.
  • ⁇ is an inflow angle with respect to the lower flow stabilizing blade 61 in parallel flow.
  • mgsin ⁇ is a component perpendicular to the lower stabilizer blade 61 of the wind force when the wind force (total amount) having the power of mg hits the lower stabilizer blade 61 at an angle ⁇ .
  • n is the n value of the height of the lower stabilizer blade 61.
  • This expression is considered to represent the following situation. That is, the wind generated from the propeller 60 hits the lower stabilizer blade 61 at the inflow angle ⁇ and flows down along the lower stabilizer blade 61 as it is, and similarly, the next wind flows down with almost no interruption. Meanwhile, since the rotation of the propeller 60 is very high, the wind on the lower stabilizer blade 61 is regarded as a continuous wind, and passes through the lower stabilizer blade 61 at a very high speed. It is considered that the force of the component perpendicular to the lower stabilizing blade 61 is applied to the lower stabilizing blade 61 instantaneously.
  • the propeller wake pk is symmetric with respect to the propeller rotation axis 65, so that the front and back surfaces of the lower stabilizer blade 61 are in opposite directions, and the component force perpendicular to the lower stabilizer blade 61 is Offset. For this reason, it is considered that the aircraft 63 did not move. However, as described above, this force is a source for stopping the anti-rotation of the airframe 63 due to the anti-torque due to the propeller rotation, and the propeller wake pk is generally said to advance in a vortex. It is.
  • Formula 5 (3) and 5- (6) Another meant to include, but not at all occur F C when inflow angle ⁇ is 0 °, once aircraft 63 begins to move, the propeller slipstream An angle is generated between the lower stabilizer blade 61 and the force of F C is applied to the airframe 63 in the direction opposite to the direction in which the airframe 63 moves, so that the higher the moving acceleration of the airframe 63 (in other words, The higher the moving speed of the fuselage 63, the larger the angle between the propeller wake and the lower stabilizing blade 61, and the magnitude of F C also increases in proportion to sin ⁇ .
  • the force F C is the sum of the components on the lower stabilizer blade 61 perpendicular to the lower stabilizer blade 61 of the total force of the propeller wake. It can be said that it becomes a source of lift commonly called.
  • the airframe 70 includes a propeller 71, a rectangular shape in a side view disposed below the propeller 71, for example, a cross-shaped lower radial stabilizer wing 72, the same height as the lower radial stabilizer wing 72, and the lower radial stabilizer
  • a cylindrical lower cylindrical stabilizing blade 73 disposed coaxially so as to surround the periphery of the blade 72 and a coaxial line disposed above the propeller 71 and having the same diameter as the lower cylindrical stabilizing blade 73.
  • a cylindrical upper cylindrical stabilizing blade 74 For example, a cylindrical upper cylindrical stabilizing blade 74, a rod-shaped connecting member 76 that connects the cylindrical stabilizing blades 73, 74, and a propeller drive unit 75 disposed on the lower radial stabilizing blade 72.
  • the diameter r 0 of the propeller 71 is assumed to be smaller than the diameter of the lower cylindrical stabilizing blade 73.
  • the wind pressure center point C due to the wake of the propeller in the lower cylindrical stabilizer wing 73 is assumed to be at a position that is 1/4 lower than the height of the lower cylindrical stabilizer wing 73 as expected. . That is, it is considered that the wind pressure center point C caused by the wake of the propeller between the lower cylindrical stabilizer 73 and the lower radial stabilizer wing 72 coincides.
  • the height of the lower cylindrical stabilizer wing 73 and the lower radial stabilizer wing 72 are both shortened, and 1 / of the diameter of the lower cylindrical stabilizer wing 73 is placed inside the lower cylindrical stabilizer wing 73.
  • the same experiment was attempted by adding an in-cylinder cylindrical stabilizer blade (not shown) having a diameter of 2. At this time, the following two assumptions 3 and 4 were made.
  • the current experiment (hereinafter referred to as the second experiment) is also the first experiment.
  • the fuselage 70 hovered quite stably.
  • the airframe 70 was rotated forward in the same direction as the rotation direction of the propeller 71, and the rotation speed was reduced by shortening the lower radial stabilizer wing 72.
  • the cylinder cylindrical stabilizing wing (diameter R 1, height h 1) to the lower tubular stable wing 73 when combining, the formula 7- (4) to formula 7- (6).
  • the shape of the stabilizing blade there are a radial shape, a cylindrical shape, an even angle regular polygonal cylindrical shape, a combination thereof, and the like.
  • a mesh shape that is symmetrical with respect to the central axis when viewed from the central axial direction.
  • the shape of the short stable wing, perpendicular to the propeller shaft may be in any shape as long as any even when viewed from a direction a shape as does not change the size of the F C.
  • a radial stabilizer blade seems to be the best.
  • the counter-torque canceling condition (hereinafter referred to as the anti-torque canceling condition) of the triangular wing airframe 63 as shown in FIG. 6 is the lower radial stabilizing wing 61 of the airframe 63 as shown in the first embodiment.
  • the n value of the height of the lower radial stabilizer 61 is given by Equation 8- (1).
  • the n value of the height of the lower radial stabilizer blade 72 is set to 1 of the value of Equation 8- (1). If it is set to 1 / 2.7 times instead of /.pi. Times, it seems that the anti-rotation of the airframe 70 can be stopped.
  • n value of the height of the lower radial stabilizer wing 61 of the triangular wing airframe 63 is also set to about 1.234, it means that the anti-rotation of the triangular wing airframe 63 also stops. This is because, in general, in the case of a triangular wing airframe, a ⁇ -fold effect appears in the force of F C when the n value of the height of the lower radial stabilizer is small.
  • the force of F C is the resistance force when the aircraft moves, it does not occur when the aircraft is stable in a vertical posture, and it works as a resistance force when the aircraft starts to tilt.
  • very small angle (inflow angle) beta is generated between the lower radial stability blades of the propeller slipstream and aircraft, F C becomes the force of only the angle amount is applied to the body it is conceivable that.
  • Equation 9- (1) is the definition equation (approximate equation) of the counter-rotation cancellation F C that best matches the previous experiments.
  • the pitch of the propeller is a distance that the propeller makes one rotation and moves forward, so the ratio of the average pitch angle (twist angle) of the two propellers with different pitches is the ratio of the above 10.95 ° and 17.82 °. It seems to appear.
  • Equation 9- (2) the n value of the height of the lower radial stabilizer blade that cancels the counter-torque due to the rotation of the propeller in the state where the ⁇ -fold effect appears (the n value at that time) The value is set as n T ), resulting in Equation 9- (2).
  • the n-value of the lower radial stabilizer for stopping counter-rotation has a width, and when the n-value is between about 1.4 to 1.6, During one hovering, the forward rotation and the reverse rotation were repeated.
  • the value of n value 1.44 is the value of n value that seems to be the most stable. Since the pitch angle at the tip of this propeller is approximately 15 °, the pitch angle on the inner side from the tip exceeds 15 °. In general lift theory, the lift decreases when the elevation angle exceeds 15 °. In the case of this propeller, the pitch angle increases more than 15 ° toward the inside of the propeller, and conversely, Lift will go down. This seems to give a wide range to the n value of the lower radial stabilizer for offsetting the propeller's counter torque.
  • n is a multiple coefficient of the propeller diameter r 0 , wind pressure (lift force) generated by the propeller wake applied to the stable wing as long as the stable wing can receive all the wake behind the propeller regardless of the spread angle of the propeller wake. ) F C increases in direct proportion to the height nr 0 of the stabilizer blade.
  • the total force ⁇ mg of one propeller rotation is applied to the stable blade in proportion to the n value.
  • the total force ( ⁇ mg) within the circle of one rotation of the propeller is almost instantaneously separated from the propeller by the wind, and flows downward without interruption. or in other words, is the same as that interruption no wind ejected from holes of diameter r 0 is given a force proportional to n times the diameter r 0 on the stability wing.
  • Wind generation of wind does not depend on the propeller (e.g. wind by explosion) is, when you are ejected from the hole of diameter r 0, if the wind pressure of the moment of wind ejected from the entire hole and P e, tubular Formula 7- (16), which is the basic formula of the wind pressure F C caused by the propeller wake on the stabilizer blade in the stabilizer blade, can be rewritten as Equation 12- (2).
  • Expressions 12- (2) to 12- (5) are expressions when the ⁇ -fold effect appears.
  • F C , F W, and W when the ⁇ -fold effect does not appear are expressed by Equations 12- (6) to 12- (9).
  • the airframe 80 in FIG. 12 has a propeller 81, a rectangular shape in a side view disposed below the propeller 81, for example, a cross-shaped lower radial stabilizer wing 82, and a coaxial line so as to surround the periphery of the lower radial stabilizer wing 82
  • a cylindrical stabilizer wing 83 whose lower end is the same height as the lower radial stabilizer wing 82 and whose upper end is extended above the propeller 81, and a propeller disposed on the lower radial stabilizer wing 82.
  • Drive unit (not shown).
  • the center of gravity G of the fuselage 80 is set to a wind pressure center point (total wind pressure center point) C of the resultant force between the wind pressure center point H caused by the propeller wake in the cylindrical stabilizer wing 83 and the wind pressure center point C caused by the propeller wake in the lower radial stabilizer wing 82. Place it at a point separated from 0 by the distance represented by Expression 7- (20), and the position of the point W is slightly larger than the distance given by Expression 7- (18), and n ⁇ 1,44. 80 was hovered.
  • the total wind pressure central point C 0 is a point obtained by dividing the distance between the points H and C by the ratio of the equation 13- (2).
  • n is the n value of the height of the lower radial stabilizer wing 82.
  • the wind pressure center point C by the propeller slipstream at lower radial stabilizing wings 82 according to the general theory of lift, and consider the case comprising an upper end a distance nr 0/4 down position of the lower radial stabilization wings 82.
  • the airframe 80 When the airframe 80 was hovered in a little wind, the airframe 80 started to be blown almost in parallel with the wind, and when the wind stopped, it was performing stable hovering at the destination. In this airframe 80, the equilibrium points of the points G and W were slightly deviated from each other, so when they were moved by the outside wind and moved in parallel, the slope was slightly adjusted but the stability as calculated.
  • the fuselage (attitude control device) 90 of FIG. 13 is a cylindrical stabilizer wing 91 and one coaxially arranged along the central axis 92 of the cylindrical stabilizer wing 91 inside the cylindrical stabilizer wing 91.
  • the radial stabilizers 93 and 94 that are symmetrical with respect to the central axis as described above (two in this example), for example, are rectangular in side view, and one or more (two in this case) that are coaxially disposed inside the cylindrical stabilizer 91.
  • each of the cylindrical stabilizing blades 91, 95, 96 is formed in a cylindrical shape, for example.
  • the diameter of the propeller 97 is assumed to be smaller than the diameter of the cylindrical stabilizing blade 91.
  • the radial stabilizer wing 93 has, for example, two stabilizer wings, and is arranged in the upper stage in the cylindrical stabilizer wing 91.
  • the radial stabilizer wing 94 has, for example, four stabilizer blades, and is disposed in the lower stage in the cylindrical stabilizer blade 91.
  • the in-cylinder cylindrical stabilizing blades 95 and 96 have different diameters, and are disposed in the lower stage in the cylindrical stabilizing blade 91 so as to intersect the lower radial stabilizing blade 94.
  • the wind pressure center points H 0 and C 1 caused by the propeller wakes in the cylindrical stabilizer wing 91 and the upper radial stabilizer wing 93 are made to coincide with each other, and the lower radial stabilizer wing 94 and each cylindrical stabilizer Wind pressure center points C 2 , H 1 , H 2 due to the propeller wake at each of the blades 95, 96 are made to coincide with each other, and the wind pressure at the wake of the propeller applied to the coincidence of the wind pressure center points H 0 , C 1 described above.
  • the distance n GC0 from the wind pressure center point C 0 to the center of gravity of the fuselage 90, the distance n GW from the external wind pressure center point W to the center of gravity of the fuselage 90, and the distance n from the fixed point 0 of the propeller rotating shaft to the center of gravity of the fuselage 90 3 elements with the G is adjusted so as to satisfy the formula 19 (5) That.
  • the following (1) to (8) can be considered as a method of adjusting the ratio of the sizes of F C1 and F C2 .
  • the position of the external wind pressure center point W can be adjusted by adding a normal blade (that is, a plate-shaped blade) to the outside of the cylindrical stabilizing blade 91.
  • the airframe 90 configured in this manner performs hovering with stability, has resistance to the influence of external wind, and can perform hovering while maintaining surprising stability.
  • the n value n 1 , n 2 of each radial stabilizer wing 93, 94 and the number m 1 of the diameter units of the stable blade of each radial stabilizer wing 93, 94, m 2 may be adjusted to satisfy Expression 14- (1).
  • ⁇ T is an inflow angle (in other words, an effective angle of the propeller 97) at which the wake of the propeller hits the main surface of each of the stabilizing blades of the radial stabilizing blades 93 and 94.
  • the n-value of the i-th radial stability blades at n i may be the number of stable wing and m i.
  • the cylindrical stabilizer wing 91 can prevent the wind flow to the radial stabilizer wings 93 and 94 from spreading outside the cylindrical stabilizer wing 91. Since the airflow can be made uniform, the stability of the aircraft 90 can be improved in the airflow.
  • in-cylinder cylindrical stabilizing blades 95 and 96 are coaxially provided inside the cylindrical stabilizing blade 91, the stability of the airframe 90 can be further improved in the wind flow.
  • the airframe 90 can be used as a propulsion device, and stable flight can be performed.
  • the radial stabilizer blades 93 and 94 and the in-cylinder cylindrical stabilizer blades 95 and 96 have the center of gravity G of the fuselage 90, the total wind pressure center point C 0 , the external wind pressure center point W, and the fixed point 0 of the propeller rotation shaft, respectively. Since it arrange
  • the counter-torque caused by the propeller rotation can be offset and the aircraft 90 can be prevented from counter-rotating due to the propeller rotation.
  • the airframe (attitude control device) 110 in FIG. 14 is a combination of two airframes 90 of S14 (hereinafter referred to as airframes 90a and 90b). More specifically, the airframe 110 has its intake-side open end directed upward, its exhaust-side open end directed downward, and its intake-side open ends inclined in opposite directions (inclination angle 0).
  • the above-mentioned two airframes 90a and 90b arranged at a distance from each other and a connecting member 111 for interconnecting the airframes 90a and 90b.
  • the connecting member 111 is formed in a substantially V-shape, for example, a rod shape bent at the center point G 0 , and a body 90 a is disposed at one end and a body 90 b is disposed at the other end. More specifically, as an extension of one end of the connecting member 111 passes through the center of gravity G 1 of and body 90a orthogonal to the central axis of the body 90a, the aircraft 90a to one end of the connecting member 111 is disposed. Similarly, as an extension of the other end of the connecting member 111 passes through the center of gravity G 2 orthogonally and body 90b to the center axis of the body 90b, aircraft 90b to the other end of the connecting member 111 is disposed.
  • the weight of each of the airframes 90a and 90b is the weight of the airframe 90a and 90b. It is desirable to maximize the rolling resistance of the fuselage 110 by increasing the height, the number of stabilizing blades, the number of stabilizing blades, and the like within the allowable range.
  • points P 1 and P 2 in FIG. 14 are the center points of the propellers 97 of the airframes 90a and 90b, respectively, and ⁇ is the line segment G 1 G 0 (and line segment G 2 G 0 ) and the horizontal direction. Is the angle between the line segment P 1 G 0 (and the line segment P 2 G 0 ) and the vertical direction, and ⁇ is the line segment G 1 G 0 and the line segment P 1. it is the angle between the angle and line G 2 G 0 and the line segment P 2 G 0 between G 0.
  • Equation 15- (4) the resultant force [F] of F and F ′ when the airframe 110 is tilted by the angle ⁇ is expressed by Equation 15- (4).
  • Equation 15- (5) is obtained from Equation 15- (3) and Equation 15- (4).
  • Expression 15- (5) can be expressed as Expression 15- (7).
  • the restoring force with respect to the shaking of the airframes 90a and 90b in the opposing direction of the airframe 110 has been described, but the restoring force with respect to the shaking of the airframe 110 in the direction perpendicular to the opposing directions of the airframes 90a and 90b.
  • the airframe uses only one device 110. It is clear that the resilience increases significantly. Therefore, it is desirable to use two or more devices 110 in combination.
  • this airframe 90c is attached to the front and rear parts and both wings of a general airplane in the direction of travel of the airplane, the stability of the airplane in the vertical and horizontal directions will be greatly increased. That is, when an airplane is flying, wind enters the aircraft 90c from the front (in the direction of the central axis of the aircraft 90c) at a high speed, and in this situation, if the aircraft 90c sways in the lateral direction (its radial direction) The wind flow from the front serves as a resistance against lateral shaking of the cylindrical stabilizer wing 91 and the stabilizer wings 93, 94, 95, 96 of the fuselage 90c, and the stability of the fuselage 90c (and thus the airplane) is greatly improved.
  • the attitude control of the aircraft or the airplane equipped with the aircraft is automatically performed by the wind pressure applied to the stable wings of the aircraft. Sensitive sensors and expensive, high-speed computer systems are unnecessary.
  • the aircraft itself or the airplane equipped with the aircraft itself is resistant to the effects of external wind, and as a result, the aircraft itself Or an airplane with that airframe will be very strong against external winds.
  • the present invention can be applied not only to propeller aircraft but also to airplanes such as jet aircraft, rockets, and gas injection, it can be widely used in fields requiring attitude control.
  • the flying unit attitude control device
  • the realization of a flying car is no longer a dream.
  • FIG. 15 is a view of a certain moment when the propeller rotates, for example, in the airframe 80 of FIG. 12 when viewed from the side.
  • point O is a fixed point of the propeller rotation axis
  • point P is an intersection of a horizontal line including the propeller operating point (center point) Y and a vertical line including point O
  • point G 1 is ,
  • the point G 2 is the position of the center of gravity when the center of gravity G of the body 80 is between the points O and P.
  • G 3 is the position of the center of gravity when the center of gravity G of the body 80 is above the point P.
  • the n value of the distance between the point O and the center of gravity G of the body 80 is n G.
  • Equation 19- (6) the moment balance equation is expressed by Equation 19- (6).
  • F C is the wind pressure behind the propeller applied to the wind pressure central point C
  • F W is the external wind pressure applied to the external wind pressure point W.
  • Equation 19- (1) is obtained in the same manner as (1) above.
  • Equation 19- (1) In order to offset the influence of the swing of the propeller rotation shaft and stabilize the airframe 80, the moment acting on the airframe 80 may be balanced, or the moment that directs the propeller rotation shaft in the vertical direction may be dominant ( That is, since the left side of Equation 19- (1) should be equal to or larger than the value of n G ), the condition for stabilizing the body 80 by offsetting the influence of the propeller rotation shaft shake is: Equation 19- (2) is obtained.
  • Equation 19- (5) the conditions for the airframe 80 to stably hover can be expressed as in Equation 19- (5).
  • Equation 19- (5) is always satisfied.
  • the total wind pressure center point of the airframe 80 is stabilized from the upper end of the cylindrical stabilizer wing 83. That is, it is fixed at a point lowered by 1/8 of the length of the wing 83.
  • the propeller position is above the point where the length of the cylindrical stabilizer wing 83 is lowered from the upper end of the cylindrical stabilizer wing 83 by one-eighth, the position of the airframe 80 is changed to the position described in S13. The total wind pressure center point appeared.
  • the position of the total wind pressure center point of the airframe 80 was obtained as follows. That is, as shown in FIG. 16, a machine body 80a in which a cylindrical auxiliary stabilizer blade 83b is attached concentrically through a connecting portion 83c below the cylindrical stabilizer blade 83 of the machine body 80 is stably hovered. By obtaining the center of gravity of the body 80a, the total wind pressure center point of the body 80 can be obtained. The ratio of the magnitudes of the total wind pressure applied to the airframe 80 and the wind pressure applied to the auxiliary stabilizing blade 83b can be calculated.
  • the position of the point at which the total wind pressure is concentrated (the total wind pressure central point) is determined. Thereafter, based on the ratio of the wind pressure applied to the entire body 80 and the wind pressure applied to the auxiliary stabilizing blade 83b, the position of the central wind pressure center point of the body 80 can be obtained.
  • the propeller position is disposed at a position lower than 1/8 or more of the length of the cylindrical stabilizer blade 83 from the upper end of the cylindrical stabilizer blade 83, the position of the radial stabilizer blade 82 can be arbitrarily selected. However, at this time, in order to stabilize the entire body 80, it is necessary to attach a cylindrical or radial auxiliary stabilizing wing 83b under the entire body 80 as shown in FIG.
  • the auxiliary stabilizer wing 83b may be attached to the auxiliary stabilizer wing 83b instead of below the fuselage 80.
  • the auxiliary stabilizer wing 83b is attached to the top, the wind pressure applied to the auxiliary stabilizer wing 83b Since it is very weak compared to the wind pressure applied, it is not realistic because the projected area of the auxiliary stabilizing blade 83b needs to be very large or mounted at a position very far from the airframe 80.
  • the total wind pressure center point C 0 of the entire fuselage 80a is the result of the wind pressure applied to the auxiliary stabilizer wing 83b. Will go down. Therefore, the position of the center of gravity for stably hovering the airframe 80a is also lowered, and the restoring effect by the center of gravity is increased. In addition, since the inertia moment of the entire body 80a is also increased, the body 80a is further stabilized as compared with the case where the auxiliary stabilizing wing 83b is not provided.
  • the means for generating the wind flow is constituted by the propeller and the propeller drive unit.
  • the present invention is not limited to this, for example, gas injection You may comprise by a machine, a jet jet machine, or a rocket jet machine. If the wind flow generating device is composed of a propeller and a propeller drive unit, it is possible to generate a wind flow on a relatively simple principle. Further, if the wind flow generating device is constituted by a gas jet, jet jet or rocket jet, a stronger wind flow can be generated.

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Abstract

Provided is a system which utilizes interaction between an airflow and a stabilizer disposed in the airflow and along the flowing direction of the airflow to secure the stability of an apparatus or an air frame integrated with the stabilizer or the stabilizer itself by the effect of the interaction. The interaction is that applying the airflow to the stabilizer at some angle changes the direction of the airflow, and the reaction of this applies force according to the reaction to the stabilizer. The stabilizer or the apparatus or the air frame integrated with the stabilizer receive the action of the force to secure stability by the effect of the action.

Description

プロペラ機、プロペラ装置および姿勢制御装置Propeller machine, propeller device and attitude control device
　本発明は、安定した垂直離着陸およびホバリングが可能なプロペラ機およびその技術の応用に関する。 The present invention relates to a propeller aircraft capable of stable vertical take-off and landing and hovering and application of the technology.
　現在において垂直離着陸および空中で静止できる機体としては、ヘリコプターおよびそれに類似した飛行機（例えばティルトローター式の米軍機Ｖ－２２）等が存在し、それ以外では、噴射口の方向を変えることのできる可変ノズルを使用した米軍機Ｆ－３５Ｂ等が存在する。これらは全て垂直離着陸時および空中静止時（ホバリング時）には、相当高度な操縦技術が要求され、その上、高度なセンサおよび高速コンピュータによる制御が不可欠になっている。このため、機体重量が増大し、製造コストも非常に大きくなり、一般機への応用は殆どできない状況である。 Currently, there are helicopters and similar airplanes (eg, tilt rotor type US military aircraft V-22), etc. as the aircraft that can be taken off and landing in the air and in the air. Otherwise, the variable direction of the injection port can be changed. There is a US military aircraft F-35B that uses a nozzle. All of these require a highly advanced maneuvering technique at the time of vertical take-off and landing and at rest in the air (at the time of hovering), and in addition, control by an advanced sensor and a high speed computer is indispensable. For this reason, the weight of the airframe increases and the manufacturing cost becomes very high, so that it can hardly be applied to general machines.
　今、この状況の中で、上記の様な問題を解決する方法が発見されれば、航空分野において飛躍的な発展が期待できる。 Now, if a method for solving the above problems is discovered in this situation, a dramatic development can be expected in the aviation field.
　尚、本発明の背景技術に関連する先行技術文献として特許文献１～８が既に公開されている。 Note that Patent Documents 1 to 8 have already been published as prior art documents related to the background art of the present invention.
実開平４－５１９９号公報Japanese Utility Model Publication No. 4-5199 特開平６－２９３２９６号公報JP-A-6-293296 特開２００６－３２７２１９号公報JP 2006-327219 A 特開２００７－１１８８９１号公報JP 2007-118891 A 特表２００５－５３３７００号公報JP 2005-533700 A 特表２００７－５２１１７４号公報Special Table 2007-521174 特開平５－３９０９２号公報JP-A-5-39092 特開平７－２３２６９９号公報Japanese Unexamined Patent Publication No. 7-232699
　これまでに開発された上述の実用機等は、自ら発した風流により不安定さが増大するという問題点があった。その上、回転翼（プロペラ）の回転振動による機体の揺れおよび横風による機体のふらつき等の他の不安定要素も加わり、垂直離着陸時およびホバリング時の機体の安定性の確保は、大変大きな問題として今も残っている。 The above-mentioned practical machines that have been developed so far have had the problem that instability increases due to the wind generated by themselves. In addition, the stability of the aircraft during vertical take-off / landing and hovering is a very big problem due to the addition of other unstable factors such as the aircraft swaying due to the rotational vibrations of the rotor blades and propeller It remains now.
　また回転翼を使用する垂直離着陸機等は回転翼の回転による反トルク相殺の方法として、二重反転翼システム、テールローターシステムおよびツインローターシステムを使用してローターの反トルクを相殺しているが、これらのシステムは構造が複雑になり、且つ操縦も大変難しいという問題点がある。 Also, vertical take-off and landing aircraft that use rotor blades counteract the rotor anti-torque by using counter-rotating blade system, tail rotor system and twin rotor system as a method of counter torque canceling by rotating rotor blades. However, these systems have a problem that the structure is complicated and the operation is very difficult.
　空を飛んでいる物体が自らの姿勢を制御する問題は、周りに自らを支える物が全く無い故に非常に難しい課題として、人類が空を飛び始めた時から人類の前に立ちはだかっていた。 The problem of controlling the posture of an object flying in the sky is a very difficult issue because there is no support for itself at all, and it has stood in front of mankind since the beginning of mankind.
　同様に従来から大規模な空調設備あるいは空洞設備等に使用するプロペラ装置では、土台への設置に際してプロペラ回転の反作用による反トルクに対して静的安定性を確保するための手段を確保するのに多きな負担が掛かるという課題があった。 Similarly, in the case of a propeller device conventionally used for large-scale air conditioning equipment or hollow equipment, etc., when installing on a base, a means for ensuring static stability against anti-torque due to the reaction of propeller rotation is secured. There was a problem that a heavy burden was applied.
　そこで、この発明の課題は、第１に、安定した垂直離着陸あるいはホバリングが可能なプロペラ機を提供すること、第２に、静的安定性を確保するための負担を低減できるプロペラ装置を提供すること、第３に、それらを応用して安定性に優れた姿勢制御装置を提供することにある。 SUMMARY OF THE INVENTION An object of the present invention is, firstly, to provide a propeller aircraft capable of stable vertical takeoff and landing or hovering, and secondly, to provide a propeller device capable of reducing a burden for ensuring static stability. Thirdly, it is to provide an attitude control device having excellent stability by applying them.
　本発明の第１の態様のプロペラ機は、前記機体は、放射状に組まれた複数の垂直主翼を備え、前記プロペラの直径をｒ0としたときに、前記プロペラが周囲の空気から受けるモーメントの２πｒ0倍のモーメントが、プロペラ後流により前記複数の垂直主翼に掛かるように、前記垂直主翼の形状が形成されるものである。 In the propeller aircraft according to the first aspect of the present invention, the aircraft includes a plurality of vertical main wings assembled radially, and when the diameter of the propeller is r 0 , the propeller has a moment received from ambient air. The shape of the vertical main wing is formed so that a moment of 2πr 0 is applied to the plurality of vertical main wings by the propeller wake.
　この態様によれば、プロペラが周囲の空気から受ける総モーメントと機体に掛かるプロペラ後流による総モーメントとを等しくでき、これにより機体におけるプロペラ回転の反作用による反回転を止める事ができる。 According to this aspect, the total moment received by the propeller from the surrounding air and the total moment caused by the propeller wake applied to the aircraft can be made equal, thereby preventing the counter-rotation due to the reaction of the propeller rotation in the aircraft.
　本発明の第２の態様のプロペラ機は、前記機体は、放射状に組まれた複数の垂直主翼を備え、前記各垂直主翼は、プロペラ後流の拡がりと同じ広さに形成され、前記各垂直主翼の高さΔｌ（Δｌ＝ｎBｒ0とする）は、前記プロペラの直径をｒ0とし、前記垂直主翼の直径単位の枚数をｍとすると、Δｌ＝２πｒ0 ／ｍを満たすように設定されるものである。 In a propeller aircraft according to a second aspect of the present invention, the airframe includes a plurality of vertical main wings that are radially assembled, and each of the vertical main wings is formed to have the same size as the spread of the propeller wake. The height of the main wing Δl (Δl = n B r 0 ) is set to satisfy Δl = 2πr 0 / m, where r 0 is the diameter of the propeller and m is the number of diameter units of the vertical main wing. It is what is done.
　この態様によれば、垂直主翼の高さΔｌをΔｌ＝２πｒ0 ／ｍを満たすように設定するだけで簡単に、機体におけるプロペラ回転の反作用による反回転を止める事ができる。 According to this aspect, the counter-rotation due to the reaction of the propeller rotation in the airframe can be stopped simply by setting the height Δl of the vertical main wing to satisfy Δl = 2πr 0 / m.
　本発明の第３の態様のプロペラ装置は、放射状に組まれた複数枚の垂直仕切板を有する機体と、前記複数枚の垂直仕切板の上側に配置されたプロペラと、を備え、前記プロペラの直径をｒ0としたときに、前記プロペラが周囲の空気から受けるモーメントの２πｒ0倍のモーメントが、プロペラ後流により前記複数の垂直仕切板に掛かるように、前記垂直仕切板の形状が形成されるものである。 A propeller device according to a third aspect of the present invention includes an airframe having a plurality of vertically partitioned vertical partition plates, and a propeller disposed on the upper side of the plurality of vertical partition plates. The shape of the vertical partition plate is formed such that when the diameter is r 0 , a moment 2πr 0 times the moment the propeller receives from the surrounding air is applied to the plurality of vertical partition plates by the wake of the propeller. Is.
　この態様によれば、プロペラが周囲の空気から受けるモーメントの２πｒ0倍のモーメントが、プロペラ後流により複数の垂直仕切板に掛かるように、垂直仕切板の形状が形成されるので、プロペラが周囲の空気から受ける総モーメントと機体に掛かるプロペラ後流による総モーメントとを等しくでき、これにより機体におけるプロペラ回転の反作用による反トルクを打ち消す事ができる。よって、静的安定性を確保するための負担を低減できる。 According to this aspect, since the shape of the vertical partition plate is formed such that a moment 2πr 0 times the moment the propeller receives from the surrounding air is applied to the plurality of vertical partition plates by the propeller wake, the propeller is The total moment received from the air and the total moment due to the propeller wake applied to the aircraft can be made equal, thereby counteracting the counter torque caused by the propeller rotation reaction in the aircraft. Therefore, the burden for ensuring static stability can be reduced.
　本発明の第４の態様のプロペラ装置は、前記各垂直仕切板は、プロペラ後流の拡がりと同じ広さに形成され、前記各垂直主翼の高さΔｌは、前記プロペラの直径をｒ0とし、前記垂直仕切板の直径単位の枚数をｍとすると、Δｌ＝２πｒ0 ／ｍを満たすように設定されるものである。 In the propeller device according to the fourth aspect of the present invention, the vertical partition plates are formed to have the same width as the spread of the propeller wake, and the height Δl of the vertical main wings is defined such that the diameter of the propeller is r 0. When the number of the vertical partition plates in the diameter unit is m, Δl = 2πr 0 / m is set.
　この態様によれば、垂直仕切板の高さΔｌをΔｌ＝２πｒ0 ／ｍを満たすように設定するだけで簡単に、機体におけるプロペラ回転の反作用による反トルクを打ち消す事ができる。 According to this aspect, the counter torque due to the reaction of the propeller rotation in the airframe can be easily canceled simply by setting the height Δl of the vertical partition plate to satisfy Δl = 2πr 0 / m.
　本発明の第５の態様のプロペラ装置は、前記機体は、前記複数枚の垂直仕切板の周囲に被された筒状体を更に有するものである。 In the propeller device according to the fifth aspect of the present invention, the airframe further includes a cylindrical body that is covered around the plurality of vertical partition plates.
　この態様によれば、プロペラ後流をプロペラの後方に集中させて噴出させることができる。 According to this aspect, the wake of the propeller can be concentrated and ejected behind the propeller.
　本発明の第６の態様のプロペラ装置は、前記機体は、ｎ（ｎ：２以上の整数）層に分割構成され、それら各層はそれぞれ、放射状に組まれた複数枚の垂直仕切板を備え、ｉ番目の前記層の高さおよび垂直仕切板の直径単位の枚数をそれぞれΔｌiおよびｍiとすると、前記各層の高さΔｌ1，Δｌ2，…，Δｌnおよび垂直仕切板の枚数ｍ1，ｍ2，，…，ｍnは、Δｌ1ｍ1＋Δｌ2ｍ2＋…＋Δｌnｍn≒２πｒ0を満たすように設定されるものである。 In the propeller device according to a sixth aspect of the present invention, the airframe is configured to be divided into n (n: an integer of 2 or more) layers, and each of the layers includes a plurality of vertical partition plates assembled radially, When i-th of the height of the layer and the number of diameter unit of a vertical partition plate respectively and .DELTA.l i and m i, height .DELTA.l 1 of each layer, Δl 2, ..., the number m 1 of .DELTA.l n and vertical partition plate , m 2 ,, ..., m n are those set so as to satisfy the Δl 1 m 1 + Δl 2 m 2 + ... + Δl n m n ≒ 2πr 0.
　この態様によれば、機体が複数層に分割構成されるので、層を追加するだけで、簡単に機体の高さを調整できる。その際、各層の高さΔｌ1，Δｌ2，，…，Δｌnおよび垂直仕切板の直径単位の枚数ｍ1，ｍ2，，…，ｍnが、Δｌ1ｍ1＋Δｌ2ｍ2＋…＋Δｌnｍn≒２πｒ0を満たすように設定されるので、機体におけるプロペラ回転の反作用による反トルクを打ち消す事ができる。 According to this aspect, since the airframe is divided into a plurality of layers, the height of the airframe can be easily adjusted simply by adding layers. At that time, the height of each layer Δl 1, Δl 2 ,, ..., number m 1, m 2 ,, ... diameter units .DELTA.l n and the vertical partition plate, m n is, Δl 1 m 1 + Δl 2 m 2 + ... Since it is set so as to satisfy + Δl n m n ≈2πr 0 , it is possible to cancel the counter torque due to the propeller rotation reaction in the airframe.
　本発明の第７の態様のプロペラ装置は、前記各層はそれぞれ、前記複数枚の垂直仕切板の周囲に被された筒状体を更に備えるものである。 In the propeller device according to the seventh aspect of the present invention, each of the layers further includes a cylindrical body covered around the plurality of vertical partition plates.
　この態様によれば、プロペラ風をプロペラ後方に有効に射出させる事ができる。 According to this aspect, the propeller wind can be effectively injected to the rear of the propeller.
　本発明の第８の態様の姿勢制御装置は、放射状安定翼または筒状安定翼を備えるものである。 The attitude control device according to the eighth aspect of the present invention includes radial stabilizers or cylindrical stabilizers.
　この態様によれば、放射状安定翼または筒状安定翼にその中心軸方向から高速度で風が進入する状況で、放射状安定翼または筒状安定翼がその径方向に揺れると、その中心軸方向からの風流が放射状安定翼または筒状安定翼の径方向の揺れに対する抵抗となり、当該姿勢制御装置（従ってこの姿勢制御装置を搭載した飛行機）の安定性を飛躍的に向上できる。 According to this aspect, when the radial stabilizer wing or the cylindrical stabilizer sways in the radial direction in the situation where the wind enters the radial stabilizer wing or the cylindrical stabilizer wing from the central axis direction at a high speed, the central stabilizer direction The wind flow from the air acts as a resistance against radial shaking of the radial stabilizer wing or the cylindrical stabilizer wing, and the stability of the attitude control device (and thus the airplane equipped with this attitude control device) can be dramatically improved.
　本発明の第９の態様の姿勢制御装置は、筒状安定翼と、前記筒状安定翼の中心軸線上に沿って同軸線状に配設された１つ以上の放射状安定翼と、を備えるものである。 An attitude control device according to a ninth aspect of the present invention includes a cylindrical stabilizer, and one or more radial stabilizers arranged coaxially along a central axis of the cylindrical stabilizer. Is.
　この態様によれば、筒状安定翼により、放射状安定翼への風流が筒状安定翼の外側に拡がる事を防止できると共にその風流を一様にできるので、風流の中において当該姿勢制御装置の安定性を向上できる。従って、この姿勢制御装置が配設された飛行機の飛行時の安定性を向上できる。 According to this aspect, the cylindrical stabilizer blade can prevent the wind flow to the radial stabilizer blade from spreading to the outside of the cylindrical stabilizer blade and can make the wind flow uniform. Stability can be improved. Therefore, the stability at the time of flight of the airplane provided with this attitude control device can be improved.
　本発明の第１０の態様の姿勢制御装置は、前記筒状安定翼と、その内部に同軸線状に１つ以上の筒内筒状安定翼を更に備えるものである。 The attitude control device according to the tenth aspect of the present invention further includes the cylindrical stabilizer and one or more in-cylinder stabilizers coaxially arranged therein.
　この態様によれば、更に風流の中において当該姿勢制御装置の安定性を向上できる。 According to this aspect, it is possible to further improve the stability of the attitude control device in the wind flow.
　本発明の第１１の態様の姿勢制御装置は、前記筒状安定翼または／および前記放射状安定翼の中心軸線上に配設され、前記中心軸線方向に風流を発生させる風流発生装置を更に備えるものである。 An attitude control device according to an eleventh aspect of the present invention further includes a wind flow generating device that is disposed on a central axis of the cylindrical stabilizer and / or the radial stabilizer and generates a wind flow in the direction of the central axis. It is.
　この態様によれば、風流発生装置を備えるので当該姿勢制御装置を推進装置として利用でき、安定した飛行を行う事ができる。 According to this aspect, since the wind flow generation device is provided, the attitude control device can be used as a propulsion device, and stable flight can be performed.
　本発明の第１２の態様の姿勢制御装置は、放射状安定翼または／および筒状安定翼は、当該姿勢制御装置の重心と各安定翼の風圧中心点の総合風圧中心点との距離ｎGCおよび重心と外部風圧中心点Ｗとの距離ｎGWとの関係が式１９-（５）で表される様に配置されるものである。 In the attitude control device according to the twelfth aspect of the present invention, the radial stabilizer wing and / or the cylindrical stabilizer wing has a distance n GC between the center of gravity of the attitude control device and the wind pressure center point of each stabilizer wing. The relationship between the center of gravity and the distance n GW between the external wind pressure center point W is arranged as represented by Expression 19- (5).
　この態様によれば、当該姿勢制御装置を安定してホバリングさせる事ができる。 According to this aspect, the attitude control device can be stably hovered.
　本発明の第１３の態様の姿勢制御装置は、前記筒状安定翼の上端から下に前記筒状安定翼の長さの１／８以上の距離の位置に、前記筒状安定翼の中心軸線方向に風流を発生させる風流発生装置を配置したものである。 The attitude control device according to the thirteenth aspect of the present invention is the central axis of the cylindrical stabilizer blade at a position of a distance of 1/8 or more of the length of the cylindrical stabilizer blade downward from the upper end of the cylindrical stabilizer blade. A wind flow generator for generating a wind flow in the direction is arranged.
　この態様によれば、当該姿勢制御装置が安定してホバリングする重心位置を変化させること無く、放射状安定翼の位置を任意に選択でき、当該姿勢制御装置の設計の自由度を向上できる。 According to this aspect, the position of the radial stabilizer can be arbitrarily selected without changing the position of the center of gravity where the attitude control device stably hovers, and the degree of freedom in designing the attitude control device can be improved.
　本発明の第１４の態様の姿勢制御装置は、前記風流発生装置が、風流を発生させるプロペラと、プロペラ用駆動部とを備えるものである。 In the attitude control device according to the fourteenth aspect of the present invention, the wind flow generator includes a propeller that generates a wind flow and a propeller drive unit.
　この形態によれば、プロペラを用いた比較的簡単な原理で風流を発生させる事ができる。 According to this embodiment, it is possible to generate a wind flow by a relatively simple principle using a propeller.
　本発明の第１５の態様の姿勢制御装置は、前記筒状安定翼の下に補助安定翼を配置させたものである。 The attitude control device according to the fifteenth aspect of the present invention is such that an auxiliary stabilizer blade is disposed under the cylindrical stabilizer blade.
　この形態によれば、当該姿勢制御装置全体の総合風圧中心点が下に下がるため、当該姿勢制御装置を安定してホバリングさせるための重心位置も下に下がることになり、重心による復元効果が増大し、より一層に安定できる。 According to this aspect, the overall wind pressure center point of the entire attitude control device is lowered, so that the position of the center of gravity for stably hovering the attitude control device is also lowered, and the restoration effect by the center of gravity is increased. However, it can be further stabilized.
　本発明の第１６の態様の姿勢制御装置は、前記放射状安定翼が中心軸線対称に形成された場合において、前記プロペラの実効角度をβTとし、ｉ番目の前記放射状安定翼の安定翼の直径単位の枚数をｍiとし、ｉ番目の前記放射状安定翼の中心軸線方向の長さを前記プロペラの直径で割った値をｎiとすると、式１４-（１）が成立するものである。 In the attitude control device according to the sixteenth aspect of the present invention, when the radial stabilizer blades are formed symmetrically with respect to the central axis, the effective angle of the propeller is β T, and the diameter of the stable blade of the i-th radial stabilizer blade is When the number of units is m i and the value obtained by dividing the length of the i-th radial stabilizer blade in the central axis direction by the diameter of the propeller is n i , Equation 14- (1) is established.
　この態様によれば、プロペラ回転による反トルクを相殺でき、当該姿勢制御装置がプロペラ回転により反回転する事を防止できる。 According to this aspect, the counter torque due to the propeller rotation can be offset, and the attitude control device can be prevented from counter rotating due to the propeller rotation.
　本発明の第１７の態様の姿勢制御装置は、第１１～第１６の何れかの姿勢制御装置のうちの同じものを２つ組み合わせた姿勢制御装置であって、それぞれその吸気側開口端を上側に向けると共にその排気側開口端を下側に向け、且つ互いの吸気側開口端を互いの対向方向に傾斜（傾斜角度０°も含む）させる様にして、互いに間隔空けて配置された前記２つの姿勢制御装置と、前記２つの姿勢制御装置を相互連結する連結部材と、を備えるものである。 A posture control device according to a seventeenth aspect of the present invention is a posture control device that is a combination of two of the eleventh to sixteenth posture control devices, each of which has its intake side open end on the upper side. And the exhaust-side opening end facing downward, and the intake-side opening ends are inclined in directions opposite to each other (including an inclination angle of 0 °). Two attitude control devices and a connecting member for interconnecting the two attitude control devices.
　この態様によれば、横揺れに対して更に安定性の増した姿勢制御装置を提供できる。 According to this aspect, it is possible to provide a posture control device with further increased stability against rolling.
実施の形態１に係るプロペラ機４０を説明するための図である。It is a figure for demonstrating the propeller machine 40 which concerns on Embodiment 1. FIG. 図１および図３のVI－VI断面図である。FIG. 4 is a sectional view taken along line VI-VI in FIGS. 1 and 3. プロペラ後流による風圧中心点の求め方を説明する図である。It is a figure explaining how to obtain | require the wind-pressure center point by the propeller wake. プロペラの一例図である。It is an example figure of a propeller. 実施の形態２に係るプロペラ装置５０Ｂの一例図である。It is an example figure of the propeller apparatus 50B which concerns on Embodiment 2. FIG. 三角翼の機体の側面図および平面図である。It is the side view and top view of a fuselage of a triangular wing. ６枚の放射状安定翼の平面視図である。It is a top view of six radial stabilizers. プロペラ回転軸の傾きをαおよびプロペラ回転軸の先端のブレの角度をβとしたときの機体の側面視概略図である。FIG. 3 is a schematic side view of the airframe when the inclination of the propeller rotation shaft is α and the angle of blur at the tip of the propeller rotation shaft is β. ｍｇの力を持つ風力が安定翼６１に作用する状態を示した図である。It is the figure which showed the state in which the wind force with the force of mg acts on the stable blade 61. FIG. プロペラ後流ｐｋが安定翼６１に作用する状態を示した図である。FIG. 6 is a diagram showing a state in which a propeller wake pk acts on a stable blade 61. 筒状安定翼および放射状安定翼を備えた機体の一例の斜視図および側面図である。It is the perspective view and side view of an example of the airframe provided with the cylindrical stabilizer blade and the radial stabilizer blade. 筒状安定翼および放射状安定翼を備えた機体の他の一例の斜視図である。It is a perspective view of another example of the airframe provided with the cylindrical stabilizer blade and the radial stabilizer blade. 筒状安定翼および放射状安定翼を備えた機体の更に他の一例の斜視図および平面図である。It is the perspective view and top view of further another example of the airframe provided with the cylindrical stabilizer blade and the radial stabilizer blade. 図１３の機体を２つ組み合わせた機体の一例図である。It is an example figure of the airframe which combined two airframes of FIG. プロペラ回転軸の振れによる影響を相殺し且つ安定するための条件を説明する図である。It is a figure explaining the conditions for canceling out and stabilizing the influence by the shake of a propeller rotating shaft. 機体８０の下方に補助安定翼８３ｂを取り付けた場合の機体８０ａの斜視図である。It is a perspective view of the airframe 80a when the auxiliary stabilizing wing 83b is attached below the airframe 80. FIG.
　＜実施の形態１＞
　この実施の形態では、図１または図３の様なプロペラ機４０Ａまたは４０Ｂのホバリングを安定させるための条件（機体が左右に揺れず、且つ機体がプロペラ回転の反作用により反回転しない条件）を検討する。 <Embodiment 1>
In this embodiment, the conditions for stabilizing the hovering of the propeller aircraft 40A or 40B as shown in FIG. 1 or FIG. 3 (conditions in which the aircraft does not swing left and right and the aircraft does not rotate due to the reaction of propeller rotation) are studied. To do.
　この実施の形態のプロペラ機４０Ａは、図１の様に、２枚の垂直主翼（以後、主翼と呼ぶ）４１ａが放射状且つ互いに平行に組まれてなる機体４１と、機体４１の上端に配置されたプロペラ４３とを備えて主構成されている。各主翼４１ａは互いに同形同大の半台形状の板状に形成されており、機体４１全体としては台形状の板状（即ちデルタ翼）に形成されている。各主翼４１ａの外側辺の傾斜角度αは、ホバリング時のプロペラ４３からの風（以後、プロペラ後流と呼ぶ）の拡がりに合わせた角度になっている。 As shown in FIG. 1, the propeller aircraft 40A of this embodiment is arranged at the upper end of an airframe 41 in which two vertical main wings (hereinafter referred to as main wings) 41a are radially and parallel to each other. And a main propeller 43. Each main wing 41a is formed in a half trapezoidal plate shape of the same shape and size, and the entire body 41 is formed in a trapezoidal plate shape (ie, a delta wing). The inclination angle α of the outer side of each main wing 41a is an angle that matches the spread of the wind from the propeller 43 during hovering (hereinafter referred to as the propeller wake).
　尚、図１中の点Ｐは、機体４１の台形を仮想的に補完してなる三角形の頂点であり、この点Ｐからプロペラ後流となる全ての風が発生していると仮定している。 It is assumed that point P in FIG. 1 is a vertex of a triangle formed by virtually complementing the trapezoid of the airframe 41, and that all the winds that follow the propeller are generated from this point P. .
　また図１中の符号Ｌは、点Ｐとプロペラ４３との間の長さであり、符号Ｒは、点Ｐと機体４１の下辺との間の長さであり、符号ｌは、プロペラ４３と機体４１の下辺との間の長さであり、符号Δｌは、機体４１の長さ（より詳細には、機体４１のプロペラ下側部分（主翼部分）の長さ）であり、符号ｒ0は、プロペラ直径であり、符号ｒlは、プロペラ４３から下に距離ｌ離れた点での機体４１の横幅（より詳細には、隣り合う２枚の主翼間での主翼表面に沿っての横幅）であり、符号ρ0 およびρlはそれぞれ、プロペラ４３直下の風量密度およびプロペラ４３から下に距離ｌ離れた点での風量密度である。 1 is the length between the point P and the propeller 43, the symbol R is the length between the point P and the lower side of the aircraft 41, and the symbol l is the length between the propeller 43 and the propeller 43. The length between the lower side of the fuselage 41 and the sign Δl is the length of the fuselage 41 (more specifically, the length of the propeller lower part (main wing part) of the fuselage 41), and the sign r 0 is The propeller diameter, and the symbol r l is the width of the airframe 41 at a distance l away from the propeller 43 (more specifically, the width along the main wing surface between two adjacent main wings) The symbols ρ 0 and ρ l are the air flow density just below the propeller 43 and the air flow density at a point away from the propeller 43 by the distance l, respectively.
　またここでは、機体４１のプロペラ下側部分（主翼部分）の高さΔｌは、プロペラ後流の最終到達距離に比べて充分に短いと想定しているので、機体４１の表面を流れるプロペラ後流の風速は、ほぼ一定とみなすことができる。 Here, it is assumed that the height Δl of the lower propeller portion (main wing portion) of the fuselage 41 is sufficiently shorter than the final reach distance of the propeller wake, so the propeller wake flowing on the surface of the fuselage 41 is assumed. The wind speed of can be regarded as almost constant.
　上記の設定の下で、プロペラ回転によってプロペラ４３が受ける反トルクを相殺する条件を求める。 条件 Under the above settings, obtain a condition for canceling the counter torque received by the propeller 43 by the propeller rotation.
　点Ｐから下に距離Ｒ離れた点でのプロペラ後流による風量密度ρは、Ｒ2 に反比例するので、式（１）となる。 Since the airflow density ρ due to the propeller wake at the point R away from the point P is inversely proportional to R 2 , Equation (1) is obtained.
Figure JPOXMLDOC01-appb-M000003
Figure JPOXMLDOC01-appb-M000003
　図１の点Ｐから全風量が下に向かって角度αで広がっている状態で、点Ｐから下に距離Ｌ離れた点の水平線上の風量密度（即ちプロペラ４３直下の風量密度）をρ0とすると、風量密度ρ0は式（２）となる。 In the state where the total air volume spreads downward from the point P in FIG. 1 by the angle α, the air volume density on the horizontal line at the point L away from the point P (that is, the air volume density directly below the propeller 43) is ρ 0. When an air volume density [rho 0 is the equation (2).
Figure JPOXMLDOC01-appb-M000004
Figure JPOXMLDOC01-appb-M000004
　Ｌtanα＝ｒ0／２だから、tanα＝ωとおくと、Ｌ＝ｒ0／２ωとなるので、式（２）は式（３）となる。 Ltanα = r 0/2 So, put the tanα = ω, since the L = r 0 / 2ω, equation (2) is the equation (3).
Figure JPOXMLDOC01-appb-M000005
Figure JPOXMLDOC01-appb-M000005
　そして式（３）を式（１）に代入すると、風量密度ρは式（４）となる。 Then, substituting equation (3) into equation (1), the air flow density ρ becomes equation (4).
Figure JPOXMLDOC01-appb-M000006
Figure JPOXMLDOC01-appb-M000006
　ここで、図１からＲ＝Ｌ＋ｌの関係が成立しているので、ｌ＝ｎｒ0（以後、このｎを距離変数と呼ぶ）とおくと、Ｒ＝（１／（２ω）＋ｎ）ｒ0となる。このＲの式を式（４）に代入すると、プロペラ４３から下に距離ｌ離れた点での風量密度ρlは、式（５）となる。 Here, since the relationship of R = L + l is established from FIG. 1, when l = nr 0 (hereinafter n is referred to as a distance variable), R = (1 / (2ω) + n) r 0 Become. Substituting this equation of R into equation (4), the air flow density ρ l at a point 1 away from the propeller 43 by the distance l becomes equation (5).
Figure JPOXMLDOC01-appb-M000007
Figure JPOXMLDOC01-appb-M000007
　そしてｒl＝２ωＲなので、式（６）となる。 Since r l = 2ωR, equation (6) is obtained.
Figure JPOXMLDOC01-appb-M000008
Figure JPOXMLDOC01-appb-M000008
　よって式（７）を得る。 Therefore, formula (7) is obtained.
Figure JPOXMLDOC01-appb-M000009
Figure JPOXMLDOC01-appb-M000009
　そして式（５）に式（７）を代入すると、式（８）を得る。 Then, substituting equation (7) into equation (5) yields equation (8).
Figure JPOXMLDOC01-appb-M000010
Figure JPOXMLDOC01-appb-M000010
　そしてプロペラ４３直下の全風量を[ρ0]とすると、[ρ0]は式（９）となる。 Then, assuming that the total air volume directly under the propeller 43 is [ρ 0 ], [ρ 0 ] is expressed by Equation (9).
Figure JPOXMLDOC01-appb-M000011
Figure JPOXMLDOC01-appb-M000011
　プロペラ４３から距離ｌ離れた点を含む水平線上の全風量を[ρl]とすると、[ρl]＝ρlｒlとなり、この[ρl]の式、式（８）および式（９）から式（１０）を得る。 If the total air volume on the horizon including the point 1 away from the propeller 43 is [ρ l ], then [ρ l ] = ρ l r l , and the equations [ρ l ], (8) and (9) ) To obtain equation (10).
Figure JPOXMLDOC01-appb-M000012
Figure JPOXMLDOC01-appb-M000012
　この実施の形態のプロペラ機４０Ｂでは、例えば図１の台形状の機体４１がプロペラ回転の反作用により反回転しないためには（即ち機体４１のプロペラ下側部分の形状を台形に保ったままで機体４１の反回転を止めるには）、どのような条件が必要かを図２および図３に基づいて検討する。 In the propeller machine 40B of this embodiment, for example, in order for the trapezoidal machine body 41 of FIG. 1 to not counter-rotate due to the reaction of the propeller rotation (that is, the machine body 41 is kept trapezoidal while maintaining the shape of the lower part of the propeller of the machine body 41). In order to stop the counter-rotation, what conditions are necessary will be examined based on FIG. 2 and FIG.
　尚、この実施の形態では、計算便宜上、機体４１は、図３の様に、プロペラ４３の上側に三角形状の安定翼４１ｃを有し、主翼４１ａと安定翼４１ｃとを合わせた形状が三角形となるように形成されている。 In this embodiment, for the sake of convenience of calculation, as shown in FIG. 3, the airframe 41 has a triangular stable wing 41c on the upper side of the propeller 43, and the combined shape of the main wing 41a and the stable wing 41c is a triangle. It is formed to become.
　図２では、プロペラ４３に当たる空気によりプロペラ４３が受ける力ａ（ａ：ベクトル）に対し、力ｂ（ｂ：ベクトル）が空気に与えられ、力ａの水平成分ａsinθの大きさの力Ｆが、機体４１のプロペラ回転軸回りの回転モーメントに関わってくる。図２で力の大きさはすべて、その力の総量を表わすとする。力ｂを受けた空気は風となり最終的に機体４１の主翼４１ａ（プロペラ下側部分）に当たり、その当たる角度θはプロペラ４３が受ける力ａの角度θとほぼ等しいとみなすことができる。よって、摩擦などの損失を無視すれば、プロペラ４３が機体４１に平行になったときのプロペラ後流による機体４１を回転させる力Ｆ’（総量）は、プロペラ４３が受ける水平成分の力Ｆ（総量）と一致し、その力Ｆ’は風の広がりを考えた風量密度の式（８）で与えられるρl（＝ρ0（ｒ0/ｒl）2）に比例すると考えられる。 In FIG. 2, a force b (b: vector) is given to the air with respect to a force a (a: vector) received by the propeller 43 by the air hitting the propeller 43, and a force F having a magnitude of the horizontal component asinθ of the force a is This is related to the rotational moment of the airframe 41 around the propeller rotation axis. In FIG. 2, all the magnitudes of the force represent the total amount of the force. The air receiving the force b becomes wind and finally hits the main wing 41a (propeller lower portion) of the fuselage 41, and the hit angle θ can be regarded as substantially equal to the angle θ of the force a received by the propeller 43. Therefore, if a loss such as friction is ignored, the force F ′ (total amount) for rotating the airframe 41 by the propeller wake when the propeller 43 is parallel to the airframe 41 is the horizontal component force F ( consistent with total), the force F 'is considered to be proportional to ([rho given by 8) l (= ρ 0 ( r 0 / r l) wherein the air volume density considering the spread of wind 2).
　言い換えれば、プロペラ４３が受ける単位面積当たりの力ｆ0およびプロペラ４３から下に任意の距離ｌ離れた点での機体４１が受ける単位面積当たりの力ｆlはそれぞれ、ｋを定数として、ｆ0＝ｋρ0、ｆl＝ｋρl と表される。 In other words, the force f 0 per unit area received by the propeller 43 and the force f l per unit area received by the airframe 41 at an arbitrary distance l downward from the propeller 43 are respectively represented by f 0. = Kρ 0 , f 1 = kρ 1
　そして、プロペラ４３が受けるプロペラ回転軸周りのモーメントＭ0は、式（１５）となる。 Then, the moment M 0 around the propeller rotation axis received by the propeller 43 is expressed by Equation (15).
Figure JPOXMLDOC01-appb-M000013
Figure JPOXMLDOC01-appb-M000013
　ここで、図３の機体４１を考えたときのプロペラ４３から下に任意の距離ｌ（＝ｎｒ0）離れた点での機体４１上の水平線上のプロペラ回転軸周りのモーメントＭlを考えると、式（１６）になる。 Here, considering the moment M l about propeller rotation axis of the horizontal line on the body 41 from the propeller 43 in a point away any distance l (= nr 0) under when considering aircraft 41 of FIG. 3 (16).
Figure JPOXMLDOC01-appb-M000014
Figure JPOXMLDOC01-appb-M000014
　式（１５）および式（１６）より式（１６ａ）が成立することが分かる。 It can be seen from equation (15) and equation (16) that equation (16a) holds.
　　Ｍ0＝Ｍl ・・・（１６ａ）
　即ち式（１６ａ）より、主翼４１ａ上の任意の距離ｌでの水平線上のモーメントは、プロペラ４３の瞬間モーメントと一致することが分かる。 M 0 = M l (16a)
That is, from the equation (16a), it can be seen that the moment on the horizontal line at an arbitrary distance l on the main wing 41a matches the instantaneous moment of the propeller 43.
　今、機体４１がホバリング状態で安定していると仮定すると、機体４１に掛かる総モーメント[Ｍl]は、機体４１のプロペラ下側部分（主翼部分）の高さをΔｌ（＝ΔｎB・ｒ0と置く）とすると、式（１７）となる。 Assuming that the fuselage 41 is stable in the hovering state, the total moment [M l ] applied to the fuselage 41 determines the height of the lower propeller portion (main wing portion) of the fuselage 41 by Δl (= Δn B · r If it is set to 0 , the following equation (17) is obtained.
Figure JPOXMLDOC01-appb-M000015
Figure JPOXMLDOC01-appb-M000015
　ここで、プロペラ４３の総モーメント[Ｍ0]を考える。プロペラ４３からの風は、現実には、プロペラ４３が機体４１にほぼ平行になった時点の風が断続的に機体４１の下に降りていくことになる。プロペラ４３が主翼４１ａと平行でない間では、プロペラ４３の各瞬間のモーメントＭ0は、主翼４１ａには伝わっていかないことになる。しかし、プロペラ４３の回転数は極めて高いため、一回転して次の平行時までの時間が非常に短く、一回転分のモーメントをすべて加えたモーメントを総モーメントと考えても良いと思われる。そうすると、プロペラ４３の総モーメント[Ｍ0]は、式（１８）になると考えられる。 Here, the total moment [M 0 ] of the propeller 43 is considered. In reality, the wind from the propeller 43 is intermittently descended under the airframe 41 when the propeller 43 becomes substantially parallel to the airframe 41. While the propeller 43 is not parallel to the main wing 41a, the moment Mo 0 of each moment of the propeller 43 is not transmitted to the main wing 41a. However, since the rotation speed of the propeller 43 is extremely high, it is considered that the time from one rotation to the next parallel time is very short, and the moment obtained by adding all the moments for one rotation may be considered as the total moment. Then, it is considered that the total moment [M 0 ] of the propeller 43 is expressed by Equation (18).
Figure JPOXMLDOC01-appb-M000016
Figure JPOXMLDOC01-appb-M000016
　そして、機体４１におけるプロペラ回転の反作用による反回転が止まる状態では、２つの総モーメント[Ｍ0]，[Ｍl]は釣り合いがとれている状態なので、式（１９）が成り立つ。 Then, in the state where the counter-rotation due to the reaction of the propeller rotation in the airframe 41 stops, the two total moments [M 0 ] and [M l ] are in a balanced state, and therefore Equation (19) is established.
Figure JPOXMLDOC01-appb-M000017
Figure JPOXMLDOC01-appb-M000017
　よって式（１７）および式（１８）よりΔｎBｒ0Ｍl＝２πｒ0Ｍ0となり、Ｍl＝Ｍ0なので、式（２０）を得る。 Therefore, from Expression (17) and Expression (18), Δn B r 0 M l = 2πr 0 M 0 and M l = M 0, so Expression (20) is obtained.
Figure JPOXMLDOC01-appb-M000018
Figure JPOXMLDOC01-appb-M000018
　式（２０）から、機体４１のプロペラ下側部分（主翼部分）の高さΔｌは、プロペラ４３の直径ｒ0の２π倍であれば、機体４１の反回転を止めることができることになる。但しこの計算は、主翼４１ａが直径単位で１枚（尚ここでは、主翼４１の枚数の数え方は直径単位で数える（即ち直径分を１枚と数える））の場合である。主翼４１ａと直交する様に更に同じく直径単位で１枚の同形同大の主翼がある場合は、式（１９）は、式（２１）となる。 From equation (20), if the height Δl of the lower propeller portion (main wing portion) of the airframe 41 is 2π times the diameter r 0 of the propeller 43, the anti-rotation of the airframe 41 can be stopped. However, this calculation is for the case where the main wing 41a is one in diameter units (here, the number of main wings 41 is counted in diameter units (that is, the diameter is counted as one)). When there is one main wing of the same shape and size in the same diameter unit so as to be orthogonal to the main wing 41a, Expression (19) becomes Expression (21).
Figure JPOXMLDOC01-appb-M000019
Figure JPOXMLDOC01-appb-M000019
　尚、式（２０）は、主翼４１ａがｍ枚ある場合は、式（４５）となる。 In addition, Formula (20) becomes Formula (45) when there are m main wings 41a.
Figure JPOXMLDOC01-appb-M000020
Figure JPOXMLDOC01-appb-M000020
　次に、図２および図３に基づいて、上記の機体４１におけるプロペラ後流に対する風圧中心点を計算する。 Next, based on FIGS. 2 and 3, the wind pressure center point for the propeller wake in the airframe 41 is calculated.
　図３の点Ｃをプロペラ後流に対する風圧中心点とし、水平線ａ，ｂ，ｃをそれぞれ機体４１のプロペラ下側部分の上辺，下辺，点Ｃを通る水平線とすると、水平線ａ，ｃ間のプロペラ後流による点Ｃを含む水平線回りのモーメントＭacと水平線ｃ，ｂ間のプロペラ後流による点Ｃを含む水平線回りのモーメントＭcbとは等しいことになる（即ち、Ｍac＝Ｍcb）。 If the point C in FIG. 3 is the wind pressure center point for the propeller wake, and the horizontal lines a, b, c are the horizontal lines passing through the upper side, the lower side, and the point C of the lower part of the propeller of the aircraft 41, respectively, the propeller between the horizontal lines a, c The moment Mac around the horizontal line including the point C due to the wake and the moment Mcb around the horizontal line including the point C due to the propeller wake between the horizontal lines c and b are equal (that is, Mac = Mcb).
　式（７）を参照して、各水平線ａ，ｂ，ｃに対応するｎ（距離変数），Ｔの値をそれぞれｎa，Ｔa，ｎb，Ｔb，ｎc，Ｔcとすると、Ｔa＝２ｎaω＋１，Ｔb＝２ｎbω＋１，Ｔc＝２ｎcω＋１の関係が成り立つ。 Referring to equation (7), if the values of n (distance variables) and T corresponding to the horizontal lines a, b and c are n a , T a , n b , T b , n c and T c , respectively, The relationship T a = 2n a ω + 1, T b = 2n b ω + 1, and T c = 2n c ω + 1 holds.
　そしてモーメントＭacは、式（６１）で表される。 The moment Mac is expressed by the equation (61).
Figure JPOXMLDOC01-appb-M000021
Figure JPOXMLDOC01-appb-M000021
　同様にモーメントＭcbは、式（６２）で表される。 Similarly, the moment Mcb is expressed by the equation (62).
Figure JPOXMLDOC01-appb-M000022
Figure JPOXMLDOC01-appb-M000022
そしてＭac＝Ｍcbから、式（６４）が成立する。 And from Mac = Mcb, Formula (64) is materialized.
Figure JPOXMLDOC01-appb-M000023
Figure JPOXMLDOC01-appb-M000023
　そして、水平線ａ，ｂに対応するＴの実測値Ｔa，Ｔbを式（６４）に代入して、プロペラ後流に対する風圧中心点Ｃに対応するＴの値Ｔcを計算し、その値Ｔcから、Ｔc＝２ｎcω＋１の関係に基づいて、Ｔcに対応する距離変数ｎの値ｎcを計算し、その値ｎcから、風圧中心点Ｃにおけるプロペラ４３からの距離ｌc（＝ｎcｒ0）を計算すれば、風圧中心点Ｃが求まる。 Then, the measured values T a and T b of T corresponding to the horizontal lines a and b are substituted into the equation (64) to calculate the value T c of T corresponding to the wind pressure center point C with respect to the wake of the propeller. from T c, based on the relationship T c = 2n c ω + 1 , to calculate the value n c of the distance variable n corresponding to T c, from the value n c, the distance l c from the propeller 43 in the wind pressure center point C By calculating (= n c r 0 ), the wind pressure center point C is obtained.
　尚、式（１８）の最右辺の「２」という数字は、プロペラ４３が２枚羽根になっている理由から２倍にしているのであるが、これは、前提としてプロペラ４３は２枚羽根を基本とし、このプロペラ４３が機体４１に平行になるときは、２枚の羽根が同時に機体４１に平行になることを想定して、機体４１全体の風量およびモーメントを計算しているので、プロペラ４３の全モーメント［Ｍ0]は、２枚の羽根の合計モーメントとしなければならないからである。ここで、羽根の枚数が増えてプロペラ４３全体のモーメント[Ｍ0]が増えても、式（４５）は変わらない。それは、プロペラ４３の枚数が増えた分、機体４１上の風量密度が大きくなるからである。従って、図４のＡ，Ｂのように、３枚羽根または４枚羽根のプロペラ４３の場合でも、式（４５）は共通して使える。 The number “2” on the rightmost side of equation (18) is doubled for the reason that the propeller 43 has two blades. However, as a premise, the propeller 43 has two blades. Basically, when the propeller 43 is parallel to the airframe 41, the airflow and moment of the entire airframe 41 are calculated on the assumption that the two blades are parallel to the airframe 41 at the same time. This is because the total moment [M 0 ] must be the total moment of the two blades. Here, even if the number of blades increases and the moment [M 0 ] of the entire propeller 43 increases, the equation (45) does not change. This is because the air flow density on the airframe 41 increases as the number of propellers 43 increases. Therefore, even in the case of a three-blade or four-blade propeller 43 as shown in FIGS.
　以上の結果に基づいた機体４１を造り、重心を上記式（６４）で計算される風圧中心点Ｃと外部風圧による中心点（以後、外部風圧中心点と呼ぶ）との間の平衡点に配置して垂直離着陸及びホバリングを行うと、理論通り、反回転も左右の揺れも全く起こらないことが実証された。尚、図３の上部安定翼４１Ｃは、外部風圧中心点の位置を調整するために設けられている。 Based on the above results, the airframe 41 is made, and the center of gravity is arranged at the equilibrium point between the wind pressure center point C calculated by the above equation (64) and the center point by the external wind pressure (hereinafter referred to as the external wind pressure center point). Then, when performing vertical take-off and landing and hovering, it was proved that neither anti-rotation nor left-right shaking occurred at all. 3 is provided for adjusting the position of the external wind pressure center point.
　以上の様に構成されたプロペラ機４０Ｂによれば、プロペラ４３が周囲の空気から受けるモーメントＭ0の２πｒ0倍のモーメント[Ｍ0]が、プロペラ後流により複数の主翼（垂直主翼）４１ａに掛かるように、垂直主翼４１ａの形状が形成されるので、プロペラが周囲の空気から受ける総モーメント[Ｍ0]と機体４１に掛かるプロペラ後流による総モーメント[Ｍl]とを等しくでき、これにより機体４１におけるプロペラ回転の反作用による反回転を止める事ができる。 According to the propeller machine 40B configured as described above, a moment [M 0 ] 2πr 0 times the moment M 0 received by the propeller 43 from the surrounding air is applied to the plurality of main wings (vertical main wings) 41a by the wake of the propeller. Since the shape of the vertical main wing 41a is formed so as to be applied, the total moment [M 0 ] received by the propeller from the surrounding air can be made equal to the total moment [M l ] due to the propeller wake applied to the airframe 41. Anti-rotation due to the reaction of propeller rotation in the airframe 41 can be stopped.
　また、各垂直主翼４１ａは、プロペラ後流の拡がりと同じ広さに形成され、各垂直主翼４１ａの高さΔｌは、Δｌ＝２πｒ0 ／ｍ（式（４５）参照）を満たすように設定されるので、垂直主翼４１ａの高さΔｌを、Δｌ＝２πｒ0 ／ｍを満たすように設定するだけで簡単に、機体４１におけるプロペラ回転の反作用による反回転を止める事ができる。 Each vertical main wing 41a is formed to have the same width as the spread of the propeller wake, and the height Δl of each vertical main wing 41a is set so as to satisfy Δl = 2πr 0 / m (see Expression (45)). Runode, the height .DELTA.l vertical wing 41a, Δl = 2πr 0 / m easily by simply set to satisfy the, can stop the counter-rotation by the reaction of the propeller rotation in the aircraft 41.
　尚、この実施の形態では、主翼４１ａの形状（主翼４１ａの傾斜角度α）は、プロペラ後流の拡がりと同じ広さに形成される場合で説明したが、プロペラ後流の拡がりよりも広く形成されても構わない。その場合は、幅ω（＝tanα）を大きくした分だけ、主翼４１ａの長さΔｌを少し短く微調整しなければならない。なぜならば、想定していたプロペラ後流の拡がりの外側にも、ある程度のプロペラ後流の流れが存在するためである。 In this embodiment, the shape of the main wing 41a (inclination angle α of the main wing 41a) has been described as being formed in the same width as the spread of the propeller wake, but it is formed wider than the spread of the propeller wake. It does not matter. In that case, the length Δl of the main wing 41a has to be slightly adjusted slightly to the extent that the width ω (= tan α) is increased. This is because there is a certain amount of propeller wake flow outside the expected spread of the propeller wake.
　尚、上記の実施の形態１では、操舵翼、制御部、駆動部および電源等については、特に記載されてないが、当然にして機体に備えられるものとする。 In the first embodiment, the steering wing, the control unit, the drive unit, the power source, and the like are not particularly described, but are naturally provided in the airframe.
　＜実施の形態２＞
　この実施の形態のプロペラ装置５０Ｂは、図５の様に、プロペラ装置５０Ｂの機体４１Ｂを複数層（ｎ層：ｎは２以上の整数。図５ではｎ＝３の場合で図示）Ｈ１，Ｈ２，…，Ｈｎに分離構成したものである。 <Embodiment 2>
As shown in FIG. 5, the propeller device 50B of this embodiment includes a plurality of layers of the airframe 41B of the propeller device 50B (n layers: n is an integer equal to or larger than 2; shown in FIG. 5 when n = 3) H1, H2 ,..., Hn.
　ｉ（ｉ＝１，２，…，ｎ）番目の層Ｈiは、放射状に組まれた複数枚ｍiの例えば矩形状の仕切板（垂直仕切板）４１ａiと、それら仕切板４１ａiの周囲に被され、それら仕切板４１ａiと同じ高さを有する例えば円筒状の筒状体４７iとを備えて構成されている。尚、筒状体４７iは、例えば仕切板４１ａiの側端面に固定されている。尚、各層Ｈi（ｉ＝１，２，…，ｎ）の仕切板４１ａiの直径単位の枚数ｍiは、同じ枚数でも、異なる枚数でも構わない。尚、各層Ｈiは、例えば各筒状体４７iの周面を介して連結部材（不図示）によって、上下に隣接するもの同士、連結固定されている。 The i (i = 1, 2,..., n) -th layer H i is composed of a plurality of radially divided m i , for example, rectangular partition plates (vertical partition plates) 41 a i and the partition plates 41 a i . For example, a cylindrical tubular body 47 i having the same height as the partition plates 41 a i is provided. Incidentally, the tubular body 47 i is, for example, fixed to the side end surface of the partition plate 41a i. The number m i of the partition units 41 a i in each layer H i (i = 1, 2,..., N) may be the same or different. Note that the layers H i are connected and fixed to each other adjacent in the vertical direction, for example, by a connecting member (not shown) via the peripheral surface of each cylindrical body 47 i .
　各層Ｈiの直径ｒiは、各層Ｈiがプロペラ回転軸４３ａに同心状に上下一列に配置されたときに、各層Ｈiの下面４７ａiがプロペラ後流の拡がりの境界線Ｑに一致する様に形成されている。各層Ｈiは、互いに間隔を空けて配置しても、互いに間隔を空けずに配置しても、どちらでも構わない。尚、各層Ｈiの下面４７ａiがプロペラ後流の拡がりの境界線Ｑから外側にはみ出す場合は、各層Ｈiの高さΔｌiを少し小さくする微調整が必要である。 The diameter r i of each layer H i, when the layers H i are arranged vertically one row concentrically propeller shaft 43a, the lower surface 47a i of each layer H i coincides with the boundary line Q spread of the propeller slipstream It is formed like this. Each layer H i is also spaced apart from one another, be arranged without an interval from each other, may either. In the case where the lower surface 47a i of each layer H i protrudes outwardly from the boundary line Q spread of the propeller slipstream is necessary fine adjustment to the slightly smaller height .DELTA.l i of each layer H i.
　そして、各層Ｈiにおけるその高さΔｌiとその仕切板４１ａiの直径単位の枚数ｍiとの積算値を、各層Ｈiに渡って足し合わせたものが、ほぼ２πｒ0となるように（即ち式（６０）を満たすように）、各層Ｈiの高さΔｌiおよび仕切板４１ａiの直径単位の枚数ｍiを設定すれば、上記の実施の形態１の場合と同様に、機体４１におけるプロペラ回転の反作用による反トルクを無くす事ができる。尚、ｒ0は、プロペラ４３の直径である。 As the integrated value of the height .DELTA.l i and the number m i of the diameter units of the partition plate 41a i in each layer H i, what the sum over each H i, becomes substantially 2.pi.r 0 ( That is, as to satisfy the equation (60)), by setting the number m i of the diameter height units .DELTA.l i and partition plates 41a i of each layer H i, as in the case of the first embodiment described above, the aircraft 41 It is possible to eliminate the counter torque due to the reaction of the propeller rotation at. R 0 is the diameter of the propeller 43.
Figure JPOXMLDOC01-appb-M000024
Figure JPOXMLDOC01-appb-M000024
　例えば図５では、ｎ＝３、ｍ1＝４、ｍ2＝３、ｍ3＝６の場合なので、この場合は、式（６０）は、４Δｌ1＋３Δｌ2＋６Δｌ3≒２πｒ0となり、この関係を満たすように各層Ｈ1，Ｈ2，Ｈ3の高さΔｌ1，Δｌ2，Δｌ3を設定すればよい。 For example, in FIG. 5, since n = 3, m 1 = 4, m 2 = 3, and m 3 = 6, in this case, equation (60) becomes 4Δl 1 + 3Δl 2 + 6Δl 3 ≈2πr 0 , and this relationship The heights Δl 1 , Δl 2 , Δl 3 of each layer H 1 , H 2 , H 3 may be set so as to satisfy the above.
　以上の様に構成されたプロペラ装置５０Ｂによれば、機体４１Ｂが複数層に分割構成されるので、層を追加するだけで、簡単に機体４１Ｂの高さ（仕切板部分の高さ）を調整できる。その際、各層の高さΔｌ1，Δｌ2，…，Δｌnおよび垂直仕切板の直径単位の枚数ｍ1，ｍ2，，…，ｍnが式（６０）を満たすように設定されるので、上記の実施の形態１の場合と同様に、機体４１Ｂにおけるプロペラ回転の反作用による反トルクを打ち消す事ができる。 According to the propeller device 50B configured as described above, since the fuselage 41B is divided into a plurality of layers, the height of the fuselage 41B (the height of the partition plate portion) can be easily adjusted simply by adding layers. it can. At that time, the height of each layer Δl 1, Δl 2, ..., the number m 1, m 2 ,, ... diameter units .DELTA.l n and the vertical partition plate, since the m n is set to satisfy equation (60) As in the case of the first embodiment, the counter torque caused by the propeller rotation reaction in the airframe 41B can be canceled.
　また各層Ｈiはそれぞれ、複数枚の垂直仕切板４１ａiの周囲に被された筒状体４７iを備えるので、プロペラ風をプロペラ後方に有効に射出させる事ができる。 The respective layers H i, so comprises a tubular body 47 i which Kabusare around the plurality of vertical partition plates 41a i, can cause the propeller wind effectively injected behind a propeller.
　＜実施の形態３＞
　この実施の形態では、下記のＳ１～Ｓ９，Ｓ１２～Ｓ１５，Ｓ１７，Ｓ１９～Ｓ２０の順に沿って、実施の形態１～２の応用例である姿勢制御装置について述べる。 <Embodiment 3>
In this embodiment, an attitude control apparatus as an application example of the first and second embodiments will be described in the order of the following S1 to S9, S12 to S15, S17, and S19 to S20.
　Ｓ１．（三角翼の機体のプロペラ後流による風圧中心点Ｃに掛かる力ＦC）
　図６の様な三角翼の機体６３を考える。機体６３は、例えば、プロペラ６０と、プロペラ６０の下側に配置された側面視台形状で放射状（例えば十字状）の下部安定翼６１と、プロペラ６０の上側に配置された側面視三角形状で放射状（例えば十字状）の上部安定翼６２と、下部放射状安定翼６１に配設されたプロペラ用の駆動部（不図示）とを備えている。 S1. (Force F C applied to wind pressure center point C by the propeller wake of the triangular wing aircraft)
Consider a triangular wing fuselage 63 as shown in FIG. The fuselage 63 has, for example, a propeller 60, a trapezoidal shape in a side view arranged on the lower side of the propeller 60, and a radial (for example, cross-shaped) lower stabilizer blade 61, and a triangular shape in a side view arranged on the upper side of the propeller 60. A radial (for example, cross-shaped) upper stabilizer blade 62 and a propeller drive unit (not shown) disposed on the lower radial stabilizer blade 61 are provided.
　尚、ｒ0、ｎCｒ0、ｎWｒ0、ｎaｒ0、ｎbｒ0を以下の様に定義し、tanα＝ωとする。 In addition, r 0 , n C r 0 , n W r 0 , n a r 0 , and n b r 0 are defined as follows, and tan α = ω.
　Ｇ：機体６０の重心
　Ｃ：プロペラ後流による風圧中心点
　Ｗ：外部風による風圧中心点
　ｒ0：プロペラ直径
　ｎaｒ0：プロペラ６０と下部安定翼６１の上辺との間の距離
　ｎbｒ0：プロペラ６０と下部安定翼６１の底辺との間の距離
　ｎcｒ0：風圧中心点Ｃのプロペラ６０からの距離
　ｎWｒ0：風圧中心点Ｗのプロペラ６０からの距離。 G: center of gravity of the aircraft 60 C: wind center point W by propeller backwash streams: external air by wind pressure center point r 0: Propeller diameter n a r 0: distance n b r between the upper side of the propeller 60 and the lower stable wing 61 0 : distance between the propeller 60 and the bottom of the lower stabilizing blade 61 n c r 0 : distance from the propeller 60 at the wind pressure center point C n W r 0 : distance from the propeller 60 at the wind pressure center point W
　また実施の形態１～２より式１-（１８）および式１-（１）が得られる。 Further, from the first and second embodiments, Formula 1- (18) and Formula 1- (1) are obtained.
Figure JPOXMLDOC01-appb-M000025
Figure JPOXMLDOC01-appb-M000025
　図６の機体６３で実験を行った。その際の機体６３の数値条件は以下である。 The experiment was conducted with the airframe 63 of FIG. The numerical conditions of the airframe 63 at that time are as follows.
　ω＝１／２、ｒ0  ＝１１．４３ｃｍ、ｎa＝０．０８、ｎb－ｎa＝π
　この数値条件の下、式１-（１８）および式１-（１）よりｎc＝１．３０４となる。 ω = 1/2, r 0 = 11.43 cm, n a = 0.08, n b −n a = π
Under this numerical condition, n c = 1.304 from Equation 1- (18) and Equation 1- (1).
　三角翼の機体６３では、外部風圧中心点Ｗのプロペラ６０からの距離のｎ値であるｎWは、機体６３の頂点Ｐから底辺までのｎ値であるπ＋１を３で割って更に２を掛けた値から、上部安定翼６２の高さのｎ値である１を引くことで求まる。よってｎW＝１．７６となる。 In the triangular wing fuselage 63, n W which is the n value of the distance from the propeller 60 of the external wind pressure center point W is divided by 3 by π + 1 which is the n value from the apex P to the base of the fuselage 63 and multiplied by two. It is obtained by subtracting 1 which is the n value of the height of the upper stabilizer blade 62 from the obtained value. Therefore, n W = 1.76.
　この機体６３を無風の中でホバリングさせて安定したホバリングを行う重心Ｇの位置を求めると、重心Ｇのｎ値であるｎGの実測値は、約１．４１９であった。ここで各点ＣＧ間の距離と各点ＧＷ間の距離との比を求めると、式１-（２）となった。 When the position of the center of gravity G at which the airframe 63 is hovered in a windless state for stable hovering is obtained, the actual measured value of n G that is the n value of the center of gravity G is about 1.419. Here, when the ratio between the distance between the points CG and the distance between the points GW is obtained, Equation 1- (2) is obtained.
Figure JPOXMLDOC01-appb-M000026
Figure JPOXMLDOC01-appb-M000026
　ここで、機体６３の下部安定翼６１の一枚分の面積ＳCと機体６３全体の投影面積（即ち上部安定翼６２の１枚分の面積と下部安定翼６１の１枚分の面積とを足した面積）ＳWとの面積比Ｗ（＝ＳW／ＳC）を計算すると、Ｗ＝１．０６であった。 Here, the area of one sheet of one sheet of the area S C and 1 sheet of the area and the lower stable wing 61 of the aircraft 63 total projected area (i.e. the upper stable wing 62 of the lower stabilizing wing 61 of the aircraft 63 When calculating the sum was area) S W and the area ratio W (= S W / S C ), were W = 1.06.
　このＷがＷ＝１のときのｎGCとｎGWとの比を求めると、式１-（２）のｎGWに係数Ｗ＝１．０６を掛ければ良いから、その結果、式１-（９）を得る。 When the W is determining the ratio of n GC and n GW when the W = 1, because it multiplied by coefficient W = 1.06 to n GW of formula 1- (2), as a result, the formula 1- ( 9) is obtained.
Figure JPOXMLDOC01-appb-M000027
Figure JPOXMLDOC01-appb-M000027
　この式１-（９）の意味するところは、プロペラ後流が流れている安定翼６１の面積ＳCに掛かるプロペラ後流による力ＦCが、同じ面積に掛かる外部風圧力ＦW’のπ倍であると言うことである。よって、機体６３全体の投影面積ＳWが、プロペラ後流が流れている安定翼６１の面積ＳCのＷ倍であるとき、一般式として式１-（１０）および式１-（１１）と表すことができる。 The meaning of Equation 1- (9) is that the force F C caused by the propeller wake acting on the area S C of the stabilizing blade 61 where the propeller wake flows is the π of the external wind pressure F W ′ applied to the same area. It is to say that it is double. Therefore, the aircraft 63 total projected area S W is, when a W times the area S C of the stable wing 61 the propeller slipstream is flowing, wherein 1- (10) as a general formula and the formula 1- (11) Can be represented.
Figure JPOXMLDOC01-appb-M000028
Figure JPOXMLDOC01-appb-M000028
　また機体６３が安定してホバリングすることから、このＦCは、機体６３の重さｍｇ（ｍ：機体６３の質量、ｇ：重力加速度）に比例した力であるはずである。よって比例係数をＫとすると、ＦC、ＦW はそれぞれ、式１-（１２）、式１-（１３）と表されると考えられる。 Further since the body 63 is hovering stable, the F C weighs mg aircraft 63 (m: mass of the aircraft 63, g: gravitational acceleration) should be a force proportional to. Therefore, when the proportionality coefficient is K, F C and F W are considered to be expressed by Expression 1- (12) and Expression 1- (13), respectively.
Figure JPOXMLDOC01-appb-M000029
Figure JPOXMLDOC01-appb-M000029
　そしてその場合のｎGCおよびｎGWの関係式は式１-（１４）となる。 In this case, the relational expression between n GC and n GW is represented by Expression 1- (14).
Figure JPOXMLDOC01-appb-M000030
Figure JPOXMLDOC01-appb-M000030
　式１-（１２）、式１-（１３）および式１-（１４）中のπは、機体６３の下部安定翼６１の高さがプロペラ６０の直径ｒ0のπ倍であることに起因していることは容易に推測できる。よってそれらの式の一般式（即ち機体６３の下部安定翼６１の高さがプロペラ６０の直径ｒ0のｎ倍である場合の式）は、式１-（１５）、式１-（１６）および式１-（１７）と表す事ができる。 Π in Formula 1- (12), Formula 1- (13), and Formula 1- (14) originates from the fact that the height of the lower stabilizer blade 61 of the fuselage 63 is π times the diameter r 0 of the propeller 60. You can easily guess what you are doing. Therefore, the general formulas of these formulas (that is, formulas in the case where the height of the lower stabilizer blade 61 of the fuselage 63 is n times the diameter r 0 of the propeller 60) are formulas 1- (15), And Formula 1- (17).
Figure JPOXMLDOC01-appb-M000031
Figure JPOXMLDOC01-appb-M000031
　機体６３の下部安定翼６１の高さのｎ値を３．０、２．９、２．８として実験を行い、式１-（１７）で表される比率の位置に重心Ｇを置くと、機体６３が安定してホバリングしたことから、式１-（１５）、式１-（１６）および式１-（１７）は、一般式と考えても良いことが分かる。 When the experiment was performed with the n value of the height of the lower stabilizer wing 61 of the airframe 63 being 3.0, 2.9, and 2.8, and the center of gravity G is placed at the position of the ratio represented by Equation 1- (17), Since the airframe 63 is stably hovered, it can be seen that Formula 1- (15), Formula 1- (16), and Formula 1- (17) may be considered as general formulas.
　Ｓ２．（下部安定翼の直径単位の枚数をｍ枚としたときの倍数係数Ｎ）
　ここでは、図６の機体６３において、図７の様に三角翼（上部安定翼６２および下部安定翼６１）を放射状に６枚設けた場合を考える。下部安定翼６１の各安定翼６１－１，６１－２，６１－３，６１－４，６１－５，６１－６に掛かるプロペラ後流による風圧力をそれぞれＦC-1，ＦC-2，ＦC-3，ＦC-4，ＦC-5，ＦC-6とする。安定翼６１－１には、それ自身に掛かるプロペラ後流による風圧力ＦC-1以外に、各安定翼６１－２～６１－６に係る風圧力ＦC-2～ＦC-6における安定翼６１－１に垂直な成分が掛かる。この事を考慮すると、安定翼６１－１に掛かる風圧力の総和力〔ＦC〕は、式２-（１）となる。 S2. (Multiplier coefficient N when the number of diameter units of the lower stabilizer blade is m)
Here, let us consider a case in which six triangular wings (upper stabilizing wing 62 and lower stabilizing wing 61) are radially provided in the fuselage 63 of FIG. 6 as shown in FIG. The wind pressures caused by the wake of the propeller applied to the stabilizing blades 61-1, 61-2, 61-3, 61-4, 61-5, 61-6 of the lower stabilizing blade 61 are respectively expressed as F C-1 and F C-2. , and F C-3, F C- 4, F C-5, F C-6. In addition to the wind pressure F C-1 caused by the wake of the propeller applied to the stabilizer blade 61-1 itself, the stabilizer blade 61-1 is stabilized at the wind pressures F C-2 to F C-6 of the stabilizer blades 61-2 to 61-6. A component perpendicular to the wing 61-1 is applied. Considering this, the total force [F C ] of the wind pressure applied to the stable blade 61-1 is expressed by Equation 2- (1).
Figure JPOXMLDOC01-appb-M000032
Figure JPOXMLDOC01-appb-M000032
　そして１-（１５）式よりＦC-1＝ＦCであり、また式２-（７）の関係式が成立するので、式２-（１）は式２-（２）と表される。 From Formula 1- (15), F C-1 = F C , and since the relational expression of Formula 2- (7) holds, Formula 2- (1) is expressed as Formula 2- (2). .
Figure JPOXMLDOC01-appb-M000033
Figure JPOXMLDOC01-appb-M000033
　ここで式２-（２）の{　}内をＮと置くと、式２-（２）は式２-（３）となる。 Here, if N is placed in {} of Equation 2- (2), Equation 2- (2) becomes Equation 2- (3).
Figure JPOXMLDOC01-appb-M000034
Figure JPOXMLDOC01-appb-M000034
　一般に下部安定翼（放射状安定翼）６１の安定翼の直径単位の枚数をｍとしたときのＮは、式２-（４）で表されることが分かる。 Generally, it can be seen that N is expressed by Equation 2- (4), where m is the diameter unit number of the stabilizing blade of the lower stabilizing blade (radial stabilizing blade) 61.
Figure JPOXMLDOC01-appb-M000035
Figure JPOXMLDOC01-appb-M000035
　尚、図７の場合、式２-（８）の関係式が成立するので、その場合のＮは、Ｎ＝２となる。 In the case of FIG. 7, since the relational expression 2- (8) is established, N in that case is N = 2.
Figure JPOXMLDOC01-appb-M000036
Figure JPOXMLDOC01-appb-M000036
　一般に下部安定翼６１が中心軸線対称で、更に下部安定翼６１の安定翼の直径単位の枚数ｍが式２-（５）で表されるとき、式２-（６）が常に成立する。 Generally, when the lower stabilizer blade 61 is symmetrical with respect to the central axis, and the number m of the stable blade diameter unit of the lower stabilizer blade 61 is expressed by Equation 2- (5), Equation 2- (6) always holds.
Figure JPOXMLDOC01-appb-M000037
Figure JPOXMLDOC01-appb-M000037
　Ｓ３．（放射状安定翼におけるプロペラ後流による風圧中心点）
　上述の様に、プロペラ後流による風圧中心点Ｃのプロペラ６０からの距離は、式１-（１）で表される。 S3. (Center point of wind pressure due to the wake of the propeller in the radial stable blade)
As described above, the distance from the propeller 60 to the wind pressure center point C due to the wake of the propeller is expressed by Expression 1- (1).
Figure JPOXMLDOC01-appb-M000038
Figure JPOXMLDOC01-appb-M000038
　この式１-（１）は、その後の実験より、ｎB（＝ｎb－ｎa）の値（下部安定翼６１の高さのｎ値）が充分に大きいときは成立するが、ｎBの値が小さくなると成立しなくなくなることが分かった。少なくともｎB≧２．６のときは、式１-（１）は成立するが、ｎB≦２．１のときは、揚力の一般理論（即ち、平行流の中の平板に対する理論）で示される位置（即ち、下部安定翼６１の上辺から下部安定翼６１の高さの１／４下がった位置）にプロペラ後流による風圧中心点Ｃが現れる。 The formula 1- (1), from subsequent experiments, n B (height of the n values of the lower stabilizing wings 61) the value of (= n b -n a) but is satisfied when sufficiently large, n B It became clear that it was not formed when the value of became small. Equation 1- (1) holds at least when n B ≧ 2.6, but when n B ≦ 2.1, it is expressed by the general theory of lift (ie, the theory for a flat plate in parallel flow). The wind pressure center point C due to the wake of the propeller appears at a position where the lower stabilizer blade 61 is lowered (ie, a position that is 1/4 lower than the height of the lower stabilizer blade 61).
　プロペラ後流による風圧中心点Ｃが、少なくともｎB＞３のときは、揚力の一般理論で表される位置ではなく式１-（１）で表される位置に移行する理由は、次の様に考えられる。揚力の一般理論は一様な平行流の中の平板に関するものであるが、プロペラ後流は一様な平行流でなくプロペラからの距離の２乗に比例して拡散するので、揚力の一般理論と一致するのは、プロペラ後流があまり拡散していない範囲だけであり、ある程度拡散した範囲では揚力の一般理論には従わなくなると思われる。 When the wind pressure center point C due to the wake of the propeller is at least n B > 3, the reason for shifting to the position represented by Formula 1- (1) instead of the position represented by the general theory of lift is as follows. Can be considered. The general theory of lift relates to a flat plate in a uniform parallel flow, but the wake behind the propeller is not a uniform parallel flow but diffuses in proportion to the square of the distance from the propeller, so the general theory of lift It is only in the range where the wake behind the propeller is not diffused much, and it seems that the general theory of lift is not followed in the range where the propeller is diffused to some extent.
　但し、機体６３を筒状にしてプロペラ後流が拡散しない様にすれば、後々の実験で実証される様に筒状の内側ではプロペラ後流は一様な平行流となり、プロペラ後流による風圧中心点Ｃは、ｎB＞３の場合でも、揚力の一般理論で表される位置に現れる。そしてここで重要なことは、後の実験で確認されるが、ＦCの一般式１-（１５）中に新たな倍数要素πが出現することである。よって式１-（１５）、式１-（１６）、式１-（１７）は、式３-（１）、式３-（２）、式３-（３）の様になる。但し、ここでＮは、前記倍数係数Ｎである。 However, if the fuselage 63 is made cylindrical so that the propeller wake does not diffuse, the propeller wake becomes a uniform parallel flow inside the cylinder as will be demonstrated in later experiments, and the wind pressure due to the propeller wake The center point C appears at a position represented by the general theory of lift even when n B > 3. What is important here is that a new multiple element π appears in the general formula 1- (15) of F C as confirmed in later experiments. Therefore, Formula 1- (15), Formula 1- (16), and Formula 1- (17) become like Formula 3- (1), Formula 3- (2), and Formula 3- (3). Here, N is the multiple coefficient N.
Figure JPOXMLDOC01-appb-M000039
Figure JPOXMLDOC01-appb-M000039
　Ｓ４．（ＦCの定義の補正）
　無風の中での図６の機体６３のホバリング実験で、プロペラ後流による風圧力ＦCが下部安定翼６１におけるプロペラ後流による風圧中心点Ｃに掛かり且つ機体６３の重心Ｇが式１-（１７）を満たさない限り、機体６３が不安定になる事実から、下記１～４が推測できる。 S4. (Correction of F C definition)
In the hovering experiment of the fuselage 63 in FIG. 6 in the absence of wind, the wind pressure F C caused by the propeller wake is applied to the wind pressure center point C caused by the propeller wake in the lower stabilizing blade 61 and the center of gravity G of the fuselage 63 is expressed by Formula 1− ( From the fact that the airframe 63 becomes unstable unless 17) is satisfied, the following 1 to 4 can be inferred.
　１．機体６３を不安定にしている原因はプロペラ後流であること
　２．ＦCなる力は、プロペラ後流が下部安定翼６１にほぼ平行に流れることから、揚力に似た力（疑似揚力）であると推測できること
　３．平行流の中の下部安定翼６１に擬似揚力が掛かるということは、平行流と思われたプロペラ後流は、実際には、下部安定翼６１に対して平行ではなく或る角度を持っていることになること
　４．プロペラ後流が下部安定翼６１に対して角度を持つ原因は、プロペラ６０の回転軸が下部安定翼６１に対して元々或る角度だけ傾いていたか、あるいはプロペラ６０の回転による回転軸の先端のブレであると考えられること。 1. 1. The reason that the airframe 63 is unstable is the wake of the propeller. 2. The force of F C can be assumed to be a force (pseudo lift) similar to lift because the wake behind the propeller flows almost parallel to the lower stabilizer blade 61. The fact that the pseudo-lift force is applied to the lower stabilizing blade 61 in the parallel flow means that the wake behind the propeller considered to be a parallel flow is actually not parallel to the lower stabilizing blade 61 but at an angle. What will happen 4. The reason why the propeller wake has an angle with respect to the lower stabilizer blade 61 is that the rotating shaft of the propeller 60 was originally inclined at a certain angle with respect to the lower stabilizer blade 61 or the tip of the rotating shaft due to the rotation of the propeller 60. Be considered to be blurring.
　これらの推測からＦCは疑似揚力であると推測して、ＦCの定義を補正する。一般に揚力Ｌの一般式の１つとして式４-（１）が知られている。 From these assumptions, F C is assumed to be a pseudo lift, and the definition of F C is corrected. In general, Formula 4- (1) is known as one of the general formulas of lift L.
Figure JPOXMLDOC01-appb-M000040
Figure JPOXMLDOC01-appb-M000040
　またＦCの定義式は式３-（１）である。 The defining formula of F C is Formula 3- (1).
Figure JPOXMLDOC01-appb-M000041
Figure JPOXMLDOC01-appb-M000041
　式４-（１）と式３-（１）の比較により、ＦCの定義式を式４-（２）の様に補正する。よってＦWの定義式も式４-（３）の様に補正される。尚、式中のｋは、疑似揚力係数であり、βは、プロペラ後流が下部安定翼（放射状安定翼）６１の各安定翼の主面に対して成す角度である。 By comparing Expression 4- (1) and Expression 3- (1), the definition of F C is corrected as shown in Expression 4- (2). Therefore, the definition formula of FW is also corrected as shown in Formula 4- (3). In the equation, k is a pseudo lift coefficient, and β is an angle formed by the propeller wake with respect to the main surface of each stabilizing blade of the lower stabilizing blade (radial stabilizing blade) 61.
Figure JPOXMLDOC01-appb-M000042
Figure JPOXMLDOC01-appb-M000042
　式４-（２）および４-（３）がほぼ正しいことは、後の実験により証明される。更にそれらの実験から、疑似揚力係数ｋはｋ≒１であることも証明される。 It is proved by later experiments that Equations 4- (2) and 4- (3) are almost correct. Furthermore, from these experiments, it is proved that the pseudo lift coefficient k is k≈1.
　Ｓ５．（擬似揚力係数ｋの決定）
　図８は、機体６３のホバリング時の不安定要素であるプロペラ回転軸の傾きをα、プロペラ回転時のプロペラ回転軸の先端のブレの角度をβとしたときの機体６３の側面視概略図である。 S5. (Determination of pseudo lift coefficient k)
FIG. 8 is a schematic side view of the airframe 63 when the inclination of the propeller rotation shaft, which is an unstable element during hovering of the airframe 63, is α, and the angle of blurring of the tip of the propeller rotation shaft during propeller rotation is β. is there.
　無風の中で機体６３がホバリングしているとすると、疑似揚力ＦCとプロペラ推進力ＦPとに関する機体６３の重力Ｇ周りのモーメントバランス式は、式５-（９）となる。 When aircraft 63 in windless is hovering, the moment balance equation about the gravity G of the pseudo-lift F C and the propeller thrust F P and related body 63 becomes Equation 5 (9).
　ここで、α、βは共に極めて小さな値とすると、式５-（９）の左辺のｍｇ［…］の項は、同式の左辺のｎGCＦCの項に対して無視できる。またＦCは、既述の通り式４-（２）で与えられる。またＦPは、推進力ＦPの鉛直成分と機体６３の重量ｍｇとの釣り合いから式５-（１０）で与えられる。 Here, if both α and β are extremely small values, the mg [...] Term on the left side of Equation 5- (9) can be ignored with respect to the n GC F C term on the left side of the equation. F C is given by the equation 4- (2) as described above. The F P is given by the formula 5- (10) from the balance between the vertical component and the weight mg of the fuselage 63 of the thrust F P.
Figure JPOXMLDOC01-appb-M000044
Figure JPOXMLDOC01-appb-M000044
　これらの事を考慮すると、式５-（９）は式５-（１）になる。 Considering these things, Equation 5- (9) becomes Equation 5- (1).
Figure JPOXMLDOC01-appb-M000045
Figure JPOXMLDOC01-appb-M000045
　この式５-（１）から疑似揚力係数ｋを求めるには、機体６３の重心Ｇと外部風圧中心点Ｗとを一致させて安定する重心Ｇの位置を決め、そのときの機体６３からｎGC，ｎC，Ｎ，ｎ，cos（α＋β）の値を求め、それらの値を式５-（１）に代入して疑似揚力係数ｋを求めればよい。 In order to obtain the pseudo lift coefficient k from the equation 5- (1), the center of gravity G of the body 63 and the external wind pressure center point W are matched to determine the position of the center of gravity G that is stable, and from the body 63 at that time, n GC , N C , N, n, cos (α + β) are obtained, and those values are substituted into the equation 5- (1) to obtain the pseudo lift coefficient k.
　尚ここでは、プロペラ回転軸の位置の固定は下部安定翼６１の上辺の中央でなされているので、式５-（１）の右辺の分子はｎCとなっているが、その固定位置を点Ｃから距離ｎXｒ0上がった位置に設定した場合は、式５-（１）のｎCはｎXに置き換えることができる。その場合の式５-（１）式（即ち一般式）は、式５-（２）になる。 In this case, since the position of the propeller rotating shaft is fixed at the center of the upper side of the lower stabilizer blade 61, the numerator on the right side of the equation 5- (1) is n C. When the position is set to a position n X r 0 higher than C , n C in Expression 5- (1) can be replaced with n X. In this case, the expression 5- (1) (that is, the general expression) becomes the expression 5- (2).
Figure JPOXMLDOC01-appb-M000046
Figure JPOXMLDOC01-appb-M000046
　無風の中、下記の条件の下で、機体６３のホバリング実験を行った（条件：sinα≒０．００６、sinβ≒０．００８、従ってcos（α＋β）≒０．９９９９≒１、下部安定翼６１の安定翼の枚数２（故にＮ＝１）、ｎ＝３．０４３６、ｎC＝１．１９４６、ｎa＝０．０９８、ｒ0＝１５．２４ｃｍ）。疑似揚力係数ｋ＝１と仮定して式５-（１）から求める距離ｎGCｒ0だけ点Ｃから下がった位置に、機体６３の重心Ｇおよび外部風圧中心点Ｗを配置して、機体６３をホバリングさせると、機体６３はほとんど安定しかける。 In the absence of wind, the hovering experiment of the airframe 63 was performed under the following conditions (conditions: sin α≈0.006, sin β≈0.008, and thus cos (α + β) ≈0.9999≈1, lower stable blade 61 The number of stable blades 2 (hence N = 1), n = 3.0436, n C = 1.1946, n a = 0.098, r 0 = 15.24 cm). A position lowered from the distance n GC r 0 only point C to obtain assuming pseudo lift coefficient k = 1 from the equation 5 (1), by placing the center of gravity G and the external wind pressure center point W of the aircraft 63, aircraft 63 When hovering, the aircraft 63 almost stabilizes.
　疑似揚力係数ｋの値を決定するこの実験では、各点Ｇ，Ｗを点Ｃから式５-（１）で求まる距離ｎGCｒ0だけ下がった位置だけでなく、その位置付近に各点Ｇ、Ｗを配置した場合についても実験を行った。その結果は、明らかに点Ｃから式５-（１）で求まる距離ｎGCｒ0だけ下がった位置に各点Ｇ，Ｗを配置した場合に限り、機体６３は安定しかかることが分かった。よって疑似揚力係数ｋの値は、ｋ＝１と考えて良いと思われる。 In this experiment for determining the value of the pseudo lift coefficient k, each point G, W is not only located at a position n GC r 0 lower than the point C by the distance n GC r 0 determined by the equation 5- (1), , W was also conducted in the experiment. As a result, it was found that the airframe 63 started to stabilize only when the points G and W were arranged at a position that was clearly lowered from the point C by the distance n GC r 0 obtained by the equation 5- (1). Therefore, it can be considered that the value of the pseudo lift coefficient k is k = 1.
　しかし、この実験で、式５-（１）で求まる距離ｎGCｒ0の位置に各点Ｇ，Ｗを配置したとき、確かに機体６３は安定しかかるが、徐々にまた横に揺れ始めることも分かった。このことは、機体６３が不安定になる原因がまだ他にあるのか、または図８で分かる様にプロペラ回転軸の先端のブレによりプロペラ後流の流入角度がα＋βとα－βとの間で揺れていることが原因なのかは分からないが、以後この疑問について実験を行った。 However, in this experiment, when the points G and W are arranged at the position of the distance n GC r 0 obtained by the equation 5- (1), the airframe 63 certainly starts to stabilize, but gradually begins to sway again. I understand. This is because there is still another cause of the instability of the fuselage 63, or as shown in FIG. 8, the inflow angle of the propeller wake is between α + β and α-β due to the shake of the tip of the propeller rotation shaft. I don't know if the cause is shaking, but I experimented on this question.
　以上の結果からＦCおよびＦWの関係式をまとめると以下の様になる。 From the above results, the relational expressions of F C and F W are summarized as follows.
　ｉ）下部安定翼６１の高さのｎ値が大きいときは（少なくともｎ≧２．６のときは）、式５-（３）、式５-（４）、式５-（５）の関係が成立する。 i) When the n value of the height of the lower stabilizer blade 61 is large (at least when n ≧ 2.6), the relationship of Equation 5- (3), Equation 5- (4), and Equation 5- (5) Is established.
Figure JPOXMLDOC01-appb-M000047
Figure JPOXMLDOC01-appb-M000047
　ｉｉ）下部安定翼６１の高さのｎ値が小さいときは、式５-（６）、式５-（７）、式５-（８）の関係が成立する。 Ii) When the n value of the height of the lower stabilizer blade 61 is small, the relationships of Equation 5- (6), Equation 5- (7), and Equation 5- (8) are established.
Figure JPOXMLDOC01-appb-M000048
Figure JPOXMLDOC01-appb-M000048
　Ｓ６．（擬似揚力ＦCの正体）
　擬似揚力ＦCが本当に一般に言われる揚力なのかどうかを確認するため、平板に掛かる揚力が下部安定翼６１に掛かっているのかを実験で確かめた。その結果は、プロペラ６０を安定翼６１に対してどの様な角度に設定しても揚力の様な力が掛かっていないことが分かった。 S6. (Identity of the pseudo-lift F C)
In order to confirm whether or not the pseudo lift F C is really a commonly known lift, it was confirmed by an experiment whether the lift applied to the flat plate is applied to the lower stabilizing blade 61. As a result, it was found that no force such as lift was applied to the propeller 60 at any angle with respect to the stable blade 61.
　今までの実験の中で何度か感じた事であるが、プロペラ６０の回転数を上げてゆき、ホバリングできる程の推進力を得た機体６３を手で横に平行移動しようとすると、プロペラ後流による風圧中心点Ｃ辺りに抵抗を感じて、機体６３はどうしても斜めに傾いてしまうことが多かった。この事からＦCなる力は抵抗力である可能性があり、それを確認する実験を行った。 As I have felt several times in the experiments so far, if I try to translate the aircraft 63 with the propulsive force enough to hover by increasing the rotation speed of the propeller 60 by hand, In many cases, the airframe 63 inevitably inclines obliquely because of the resistance around the wind pressure center point C due to the wake. For this reason, there is a possibility that the force of F C may be a resistance force, and an experiment was conducted to confirm it.
　ＦCなる力が抵抗力であるならば、上記のＳ５の実験において、機体６３の重心Ｇと風圧中心点Ｃとの位置を逆にすると（即ち重心Ｇを風圧中心点Ｃよりも上に配置すると）、それに伴ってＦCの向きは逆向きになるはずである。そしてその場合のモーメントバランス式から式５-（２）に対応する式を求め、その式を用いて、同じ条件の下で同じホバリング実験を行えば、同じ結果（即ち機体６３が安定しかかるという結果）が得られるはずである。 If the force F C is a resistance force, the position of the center of gravity G of the airframe 63 and the wind pressure center point C are reversed in the experiment of S5 described above (that is, the center of gravity G is disposed above the wind pressure center point C). then), the orientation of the F C should be reversed accordingly. Then, if an equation corresponding to Equation 5- (2) is obtained from the moment balance equation in that case, and the same hovering experiment is performed under the same condition using the equation, the same result (ie, the airframe 63 will be stabilized). Result) should be obtained.
　そこで実際に、重心Ｇと風圧中心点Ｃとの位置を逆にし（即ち重心Ｇを風圧中心点Ｃより上に配置し）、更にＦCの向きを逆にしたモーメントバランス式を考えると、式５-（２）に対応する式として式６-（１）が得られる。ここで、ｋ＝１としたときの式５-（２），式６-（１）を書き換えると、式６-（２），式６-（３）となる。 Therefore, when considering the moment balance equation in which the positions of the center of gravity G and the wind pressure center point C are actually reversed (that is, the center of gravity G is disposed above the wind pressure center point C) and the direction of F C is further reversed. Equation 6- (1) is obtained as an equation corresponding to 5- (2). Here, when Equations 5- (2) and 6- (1) when k = 1 are rewritten, Equations 6- (2) and Equation 6- (3) are obtained.
Figure JPOXMLDOC01-appb-M000049
Figure JPOXMLDOC01-appb-M000049
　そしてこの式６-（３）で求まる距離ｎGCｒ0だけ風圧中心点Ｃよりも上に各点Ｇ，Ｗを配置して上記のＳ５の実験と同じ実験を行うと、その結果は、上記のＳ５の実験結果と全く同じであった。 When the points G and W are arranged above the wind pressure center point C by the distance n GC r 0 obtained by the equation 6- (3) and the same experiment as the experiment of S5 is performed, the result is as follows. The result was exactly the same as that of S5.
　以上より、ＦCなる力は機体６３がどちらの方向に動いても、どのような方向に傾きかけても発生する抵抗力であると思われる。この事が、前記実験で安定しかかった機体がまた不安定になりかける原因ではないかと考えられる。そこで、この実験のｎGCに対応したｎGWを式５-（５）から求め、そのｎGWで決まる距離ｎGWｒ0の位置に点Ｗを配置して、再びホバリング実験を行うと、機体６３は安定してホバリングした。 From the above, it can be considered that the force F C is a resistance force that is generated regardless of the direction in which the airframe 63 moves, regardless of the direction. This is considered to be the reason why the aircraft that has been stabilized in the experiment is becoming unstable again. Therefore, the n GW corresponding to the n GC of this experiment is obtained from the equation 5- (5), the point W is arranged at the position of the distance n GW r 0 determined by the n GW , and the hovering experiment is performed again. 63 stably hovered.
　これらの実験による検証により、ＦCなる力は機体６３が動こうとしたときの抵抗力であることが分かった。よって機体６３が全く動いていないときには何らの力も発生しないことが分かる。ＦCなる力が抵抗力であるならば、ＦCは外部風に対する抵抗となり、ホバリング時の外部風に対する機体６３の安定性が増大することになる。外部風により機体６３が押されて動き始めたときにＦCなる抵抗力が発生し、その動きの加速度が低減されることとなる。 From the verification by these experiments, it was found that the force of F C is a resistance force when the airframe 63 tries to move. Therefore, it can be seen that no force is generated when the airframe 63 is not moving at all. If F C becomes the force is resistant, F C becomes a resistance to external air, so that the stability of the machine body 63 is increased with respect to the external air during hovering. When the airframe 63 is pushed by the external wind and starts to move, a resistance force F C is generated, and the acceleration of the movement is reduced.
　ここでＦCの定義式である式５-（６）の意味を考える。 Consider the meaning of Formula 5- (6), which is the definition formula of F C.
Figure JPOXMLDOC01-appb-M000050
Figure JPOXMLDOC01-appb-M000050
　式５-（６）中のβは、平行流の下部安定翼６１に対する流入角度である。図９の様に、ｍｇsinβは、ｍｇの力を持つ風力（総量）が下部安定翼６１に角度βで当たった際のその風力の下部安定翼６１に垂直な成分である。ｎは、下部安定翼６１の高さのｎ値である。Ｎは、下部安定翼６１の安定翼の直径単位の枚数による倍数係数である。説明便宜上、安定翼の直径単位の枚数が２枚（即ちＮ＝１）の場合を考える。まずｎ値が大きい場合（即ちＦCが式５-（３）で表される場合）を考える。即ちこの場合のＦCは式５-（３）’となる。 In Expression 5- (6), β is an inflow angle with respect to the lower flow stabilizing blade 61 in parallel flow. As shown in FIG. 9, mgsin β is a component perpendicular to the lower stabilizer blade 61 of the wind force when the wind force (total amount) having the power of mg hits the lower stabilizer blade 61 at an angle β. n is the n value of the height of the lower stabilizer blade 61. N is a multiple coefficient according to the number of diameter units of the stabilizing blades of the lower stabilizing blade 61. For convenience of explanation, consider the case where the number of stabilizing blades is 2 (ie, N = 1). First, consider the case where the n value is large (that is, the case where F C is expressed by Equation 5- (3)). That is, F C in this case is expressed by Equation 5- (3) ′.
Figure JPOXMLDOC01-appb-M000051
Figure JPOXMLDOC01-appb-M000051
　この式は、以下の状況を表していると考えられる。即ち、プロペラ６０から発せられた風が下部安定翼６１に流入角度βで当たりそのまま下部安定翼６１に沿って下に流れて行き、同様にほとんど間断無く次の風が当たり下に流れて行く。その間、プロペラ６０の回転が非常に高速なので下部安定翼６１上の風は連続した風とみなされ、また非常に高速で下部安定翼６１を通り過ぎるから、下部安定翼６１上の全ての風力（総量）の総和の下部安定翼６１に垂直な成分の力が下部安定翼６１に一瞬に掛かっていると考えられる。 This expression is considered to represent the following situation. That is, the wind generated from the propeller 60 hits the lower stabilizer blade 61 at the inflow angle β and flows down along the lower stabilizer blade 61 as it is, and similarly, the next wind flows down with almost no interruption. Meanwhile, since the rotation of the propeller 60 is very high, the wind on the lower stabilizer blade 61 is regarded as a continuous wind, and passes through the lower stabilizer blade 61 at a very high speed. It is considered that the force of the component perpendicular to the lower stabilizing blade 61 is applied to the lower stabilizing blade 61 instantaneously.
　では、下部安定翼６１の高さのｎ値が小さい場合に、ＦCの定義式に係数πが現れるのはなぜなのか。その様な場合は風量密度が大きくなり、プロペラ６０の１回転の力の総和πｍｇの下部安定翼６１に垂直な成分の総和がｎπｍｇsinβになると考えられる。しかしこのπ倍になる効果（π倍効果）が実際なぜ起こるのかはわからない。 Then, why does the coefficient π appear in the definition formula of F C when the n value of the height of the lower stabilizer blade 61 is small? In such a case, it is considered that the air flow density increases, and the sum of the components perpendicular to the lower stable blade 61 of the total rotation force πmg of the propeller 60 becomes nπmgsinβ. However, we do not know why this effect of π (pi-fold effect) actually occurs.
　ＦCの定義式では、流入角度βが少しでもあれば疑似揚力ＦCが発生し、機体６３はプロペラ６０の傾いた方向とは逆方向に動き始めるはずである。しかし実験では、プロペラ６０を下部安定翼６１に対してどのような角度に設定しても、疑似揚力ＦCは発生しなかったのはなぜなのか。 In the definition formula of F C , if the inflow angle β is small, a pseudo lift F C is generated, and the fuselage 63 should start to move in the direction opposite to the direction in which the propeller 60 is inclined. However, in experiments, be set at any angle to the propeller 60 to the lower stable wing 61, pseudo-lift F C is why thing is did not occur.
　図１０を参照してその原因と思われる理由を説明する。図１０から明らかな様に、プロペラ後流ｐｋは、プロペラ回転軸６５に対して対称なので、下部安定翼６１の表面と裏面とでは逆方向となり、下部安定翼６１に垂直なその成分の力は相殺される。このため機体６３は動かなかったものと考えられる。しかしこの力は、既述の通り、プロペラ回転による反トルクによる機体６３の反回転を止めるための源泉となっており、一般にプロペラ後流ｐｋは渦を巻いて進んで行くと言われている所以である。 Referring to FIG. 10, the reason that seems to be the cause will be described. As is clear from FIG. 10, the propeller wake pk is symmetric with respect to the propeller rotation axis 65, so that the front and back surfaces of the lower stabilizer blade 61 are in opposite directions, and the component force perpendicular to the lower stabilizer blade 61 is Offset. For this reason, it is considered that the aircraft 63 did not move. However, as described above, this force is a source for stopping the anti-rotation of the airframe 63 due to the anti-torque due to the propeller rotation, and the propeller wake pk is generally said to advance in a vortex. It is.
　この事からプロペラ後流では揚力（疑似揚力）は発生しないが、揚力の源泉となる風力は存在することが分かる。 From this, it can be seen that there is no lift (pseudo lift) in the wake of the propeller, but there is wind power that is the source of lift.
　式５-（３）および式５-（６）のもう１つの意味することは、流入角度βが０°のときはＦCは全く発生しないが、ひとたび機体６３が動き始めると、プロペラ後流と下部安定翼６１の間に角度が発生し、機体６３の動く方向とは逆方向にＦCなる力が機体６３に掛かるということで、機体６３の動く加速度が大きければ大きい程（換言すれば、機体６３の動く速度が大きければ大きい程）プロペラ後流と下部安定翼６１の角度が大きくなり、ＦCの大きさもsinβに比例して大きくなるということである。 Formula 5 (3) and 5- (6) Another meant to include, but not at all occur F C when inflow angle β is 0 °, once aircraft 63 begins to move, the propeller slipstream An angle is generated between the lower stabilizer blade 61 and the force of F C is applied to the airframe 63 in the direction opposite to the direction in which the airframe 63 moves, so that the higher the moving acceleration of the airframe 63 (in other words, The higher the moving speed of the fuselage 63, the larger the angle between the propeller wake and the lower stabilizing blade 61, and the magnitude of F C also increases in proportion to sin β.
　以上の事からＦCなる力は、プロペラ後流の総量の力の下部安定翼６１に垂直な成分の下部安定翼６１上の総和であると言うことができる。一般に言われる揚力の源泉になると言える。 From the above, it can be said that the force F C is the sum of the components on the lower stabilizer blade 61 perpendicular to the lower stabilizer blade 61 of the total force of the propeller wake. It can be said that it becomes a source of lift commonly called.
　Ｓ７．（筒状安定翼）
　図１１の様な機体７０を考える。この機体７０は、プロペラ７１と、プロペラ７１の下側に配置された側面視矩形状で例えば十字状の下部放射状安定翼７２と、その下部放射状安定翼７２と同じ高さで且つその下部放射状安定翼７２の周囲を囲繞する様に同軸線状に配設された例えば円筒状の下部筒状安定翼７３と、プロペラ７１の上側に同軸線状に配置され、下部筒状安定翼７３と同径の例えば円筒状の上部筒状安定翼７４と、各筒状安定翼７３，７４を相互に連結する例えば棒状の連結部材７６と、下部放射状安定翼７２に配設されたプロペラ用の駆動部７５とを備える。プロペラ７１の直径ｒ0は、下部筒状安定翼７３の口径よりも小さいものとする。 S7. (Cylindrical stabilizer)
Consider an aircraft 70 as shown in FIG. The airframe 70 includes a propeller 71, a rectangular shape in a side view disposed below the propeller 71, for example, a cross-shaped lower radial stabilizer wing 72, the same height as the lower radial stabilizer wing 72, and the lower radial stabilizer For example, a cylindrical lower cylindrical stabilizing blade 73 disposed coaxially so as to surround the periphery of the blade 72 and a coaxial line disposed above the propeller 71 and having the same diameter as the lower cylindrical stabilizing blade 73. For example, a cylindrical upper cylindrical stabilizing blade 74, a rod-shaped connecting member 76 that connects the cylindrical stabilizing blades 73, 74, and a propeller drive unit 75 disposed on the lower radial stabilizing blade 72. With. The diameter r 0 of the propeller 71 is assumed to be smaller than the diameter of the lower cylindrical stabilizing blade 73.
　この機体７０において、外部風圧中心点Ｗの位置を調整しながら機体７０の揺れが安定する機体７０の重心Ｇの位置を求め、その際のｎGCとｎGWとの比がどの様になるかを測定した。プロペラ後流は下部筒状安定翼７３の周りには全く漏れていないことを確認して、無風の中で実験を行った。 In this airframe 70, the position of the center of gravity G of the airframe 70 where the shaking of the airframe 70 is stabilized is obtained while adjusting the position of the external wind pressure center point W, and what is the ratio of nGC and nGW at that time? Was measured. The experiment was conducted in a windless state after confirming that the propeller wake did not leak around the lower cylindrical stabilizer 73 at all.
　下部筒状安定翼７３におけるプロペラ後流による風圧中心点Ｃは、予想として下部筒状安定翼７３の上面からその下部筒状安定翼７３の高さの１／４下がった位置にあると思われる。即ち下部筒状安定翼７３と下部放射状安定翼７２とのプロペラ後流による風圧中心点Ｃは一致すると思われる。 The wind pressure center point C due to the wake of the propeller in the lower cylindrical stabilizer wing 73 is assumed to be at a position that is 1/4 lower than the height of the lower cylindrical stabilizer wing 73 as expected. . That is, it is considered that the wind pressure center point C caused by the wake of the propeller between the lower cylindrical stabilizer 73 and the lower radial stabilizer wing 72 coincides.
　尚、実験の各条件値は、ｎ＝π、Ｎ＝１、ｎX＝０．７１３４，ｒ0＝１５．２４ｃｍである。 The condition values of the experiment are n = π, N = 1, n X = 0.7134, r 0 = 15.24 cm.
　またここで、下記の２つの仮定１，２を置き、式６-（２）および式５-（８）を用いてｎGCおよびｎGWの値を求め、それらの値から求まる重心および外部風圧中心点に実際の重心Ｇおよび外部風圧中心点Ｗを配置して実験を試みた。 Also, here, the following two assumptions 1 and 2 are set, and the values of n GC and n GW are obtained using equations 6- (2) and 5- (8), and the center of gravity and external wind pressure obtained from these values are obtained. An experiment was attempted by placing the actual center of gravity G and the external wind pressure center point W at the center point.
　尚、上記の２つの仮定１，２は、以下の通りである。 The above two assumptions 1 and 2 are as follows.
　仮定１：下部筒状安定翼７３に掛かるプロペラ後流による風圧力ＦCは、下部放射状安定翼７２の安定翼１枚分（即ち下部筒状安定翼７３の高さ×直径の面積）に掛かるＦCと同じである。 Assumption 1: The wind pressure F C due to the propeller wake applied to the lower cylindrical stabilizing blade 73 is applied to one stabilizing blade of the lower radial stabilizing blade 72 (that is, the height of the lower cylindrical stabilizing blade 73 × the area of the diameter). Same as F C.
　仮定２：面積比Ｗ（＝ＳW／ＳC）を求めるときの面積ＳCは、下部放射状安定翼７２の安定翼１枚分だけではなく、更に仮定１を考慮して下部筒状安定翼７の安定翼１枚分の面積も加えた値とする。換言すれば、下部放射状安定翼７２の安定翼１枚の面積の２倍の面積をＳCとする。 Assumption 2: the area S C when obtaining the area ratio W (= S W / S C ) , the lower radial stability blade 72 not only stable wing one sheet of, further lower tubular stable wing in consideration of the assumptions 1 The area for one stable blade of 7 is also added. In other words, twice the area of the stabilizing wings one area of the lower radial stabilization wings 72 and S C.
　この仮定に基づき式６-（２）および式５-（８）を用いてｎGCおよびｎGWを計算すると、ｎGC＝０．０３８（従ってｎGCｒ0＝０．５８ｃｍ）およびｎGW＝１．００（従ってｎGWｒ0＝１５．２４ｃｍ）となった。そしてこの計算値ｎGWに一致する様に、機体７０の上部筒状安定翼７４と下部筒状安定翼７３との間隔ｘｒ0のｘ値を計算すると、ｘ＝２．８７５（従ってｘｒ0＝４３，８ｃｍ）となった。 Based on this assumption, n GC and n GW are calculated using Equation 6- (2) and Equation 5- (8), and n GC = 0.038 (hence n GC r 0 = 0.58 cm) and n GW = 1.00 (hence n GW r 0 = 15.24 cm). When the x value of the distance xr 0 between the upper cylindrical stabilizer wing 74 and the lower cylindrical stabilizer wing 73 of the fuselage 70 is calculated so as to coincide with the calculated value n GW , x = 2.875 (accordingly, xr 0 = 43.8 cm).
　上記の条件値および計算値を満たす様に機体７０を調整すると、機体７０は安定してホバリングした。 When the aircraft 70 was adjusted to satisfy the above condition values and calculated values, the aircraft 70 was stably hovered.
　図１１の機体７０の実験では、左右の揺れは全く無く安定してホバリングを行ったが、機体７０はプロペラ７１の回転方向と同じ方向に正回転していた。この正回転を止めるためには、下部放射状安定翼７２の高さを短くすれば良いことは明らかである。この正回転を止める放射状安定翼７２の高さは、後の実験で明らかになる。 In the experiment of the airframe 70 in FIG. 11, there was no left / right shaking and the hovering was performed stably, but the airframe 70 was rotating forward in the same direction as the rotation direction of the propeller 71. Obviously, in order to stop this forward rotation, the height of the lower radial stabilizer blade 72 may be shortened. The height of the radial stabilizing blade 72 that stops the forward rotation will become clear in later experiments.
　次に図１１の機体７０において、下部筒状安定翼７３および下部放射状安定翼７２の高さを共に短くして、下部筒状安定翼７３の内部に下部筒状安定翼７３の直径の１／２の直径の筒内筒状安定翼（不図示）を追加して同じ実験を試みた。このときも下記の２つの仮定３，４を置いた。 Next, in the fuselage 70 of FIG. 11, the height of the lower cylindrical stabilizer wing 73 and the lower radial stabilizer wing 72 are both shortened, and 1 / of the diameter of the lower cylindrical stabilizer wing 73 is placed inside the lower cylindrical stabilizer wing 73. The same experiment was attempted by adding an in-cylinder cylindrical stabilizer blade (not shown) having a diameter of 2. At this time, the following two assumptions 3 and 4 were made.
　仮定３：筒内筒状安定翼の側壁には、外側の筒状安定翼７３の側壁の様に内側だけなく内外両側にプロペラ後流が流れるので、筒内筒状安定翼に掛かるプロペラ後流による風圧力ＦCは、下部放射状安定翼７２の幅１／２の安定翼２枚分に掛かるＦCと同じである。 Assumption 3: Since the propeller wake flows not only on the inner side but also on both the inner and outer sides as in the side wall of the outer cylindrical stabilizer wing 73 on the side wall of the cylindrical cylindrical stabilizer wing, the wake of the propeller on the cylindrical stabilizer wing The wind pressure F C due to is the same as the F C applied to two ½ width stable blades of the lower radial stabilizer wing 72.
　仮定４：面積比Ｗ（＝ＳW／ＳC）を求めるときには、下部放射状安定翼７２の安定翼１枚分の面積を既述の仮定１を考慮して２倍にした面積に、更に仮定３を考慮して下部放射状安定翼７２の安定翼１枚分の面積を加えた面積をＳCとする。換言すれば、下部放射状安定翼７２の安定翼１枚の面積の３倍の面積をＳCとする。 Assumption 4: When obtaining the area ratio W (= S W / S C ), further assume that the area of one stabilizing blade of the lower radial stabilizing blade 72 is doubled in consideration of the aforementioned assumption 1. the area obtained by adding the stabilizing wing area of one sheet of the lower radial stability blade 72 in view of the 3 and S C. In other words, three times the area of the stabilizing wings one area of the lower radial stabilization wings 72 and S C.
　以上の仮定に基づき図１１の機体７０での実験（以後、第１の実験と呼ぶ）と同じ実験を行うと、今回の実験（以後、第２の実験と呼ぶ）でも、第１の実験と同様に、機体７０は全く安定してホバリングをした。しかし今回の実験でも、第１の実験と同様に、機体７０はプロペラ７１の回転方向と同方向に正回転し、その回転速度は、下部放射状安定翼７２を短くした分、小さくなっていた。 Based on the above assumptions, if the same experiment as that of the airframe 70 in FIG. 11 (hereinafter referred to as the first experiment) is performed, the current experiment (hereinafter referred to as the second experiment) is also the first experiment. Similarly, the fuselage 70 hovered quite stably. However, in this experiment as well, as in the first experiment, the airframe 70 was rotated forward in the same direction as the rotation direction of the propeller 71, and the rotation speed was reduced by shortening the lower radial stabilizer wing 72.
　下部筒状安定翼７３および上部筒状安定翼７４の各々の直径および下部放射状安定翼７２の横幅を共に１．１１５ｒ0（ｒ0＝１５．２４ｃｍ）と大きくして上記の第１の実験および第２の実験を再度行った。それらの実験結果が同じであったことから、下部筒状安定翼７３の外側にプロペラ後流が漏れない様にすれば、下部筒状安定翼７３の直径および下部放射状安定翼７２の横幅に関係なく、同じ結果（安定してホバリングするという結果）が得られることが分かった。 The first experiment and the by increasing the both 1.115R 0 the width of each of the diameter and the lower radial stabilization wings 72 of the lower tubular stable wing 73 and the upper tubular stable wing 74 (r 0 = 15.24cm) A second experiment was performed again. Since the experimental results were the same, if the propeller wake was not leaked outside the lower cylindrical stabilizer 73, it was related to the diameter of the lower cylindrical stabilizer 73 and the lateral width of the lower radial stabilizer 72. It was found that the same result (result of stable hovering) was obtained.
　上記の第１および第２の実験により、下部筒状安定翼（即ち外側の筒状安定翼）７３に対するＦC、ＦW、Ｗ、ｎGC、ｎGWの各関係式は次の様になる。下部筒状安定翼７３の高さのｎ値をｈ0とし、下部筒状安定翼７３の直径をＲ0とすると式７-（１）～式７-（３）となる。 From the first and second experiments, the relational expressions of F C , F W , W, n GC , and n GW for the lower cylindrical stabilizer (ie, the outer cylindrical stabilizer) 73 are as follows. . If the n value of the height of the lower cylindrical stabilizing blade 73 is h 0 and the diameter of the lower cylindrical stabilizing blade 73 is R 0 , Equations 7- (1) to 7- (3) are obtained.
Figure JPOXMLDOC01-appb-M000052
Figure JPOXMLDOC01-appb-M000052
　またこの下部筒状安定翼７３内に筒内筒状安定翼（直径Ｒ1，高さｈ1）を組み合わせた場合は、式７-（４）～式７-（６）となる。 The cylinder cylindrical stabilizing wing (diameter R 1, height h 1) to the lower tubular stable wing 73 when combining, the formula 7- (4) to formula 7- (6).
Figure JPOXMLDOC01-appb-M000053
Figure JPOXMLDOC01-appb-M000053
　更に下部筒状安定翼７３、筒内筒状安定翼および放射状安定翼７２を組み合わせた場合は、式７-（７）～式７-（９）となる。 Further, when the lower cylindrical stabilizer wing 73, the in-cylinder cylindrical stabilizer wing, and the radial stabilizer wing 72 are combined, Expressions 7- (7) to 7- (9) are obtained.
Figure JPOXMLDOC01-appb-M000054
Figure JPOXMLDOC01-appb-M000054
　これらの式は全て、下部筒状安定翼７３、筒内筒状安定翼および放射状安定翼７２の各々のプロペラ後流による風圧中心点Ｃを一致させたときの合力ＦCに対する式である。 All of these expressions are equations for force F C when to match the wind pressure center point C by respective propeller slipstream of lower tubular stable wing 73, cylinder cylindrical stabilizing wings and radial stability blades 72.
　ｍ個の筒内筒状安定翼を組み合わせたときには、それぞれの筒内筒状安定翼の直径Ｒmの下部筒状安定翼の直径Ｒ0に対する比Ｒm／Ｒ0をεmとすれば、式７-（７）～式７-（９）は、一般式として式７-（１０）～式７-（１２）となる。 When m in-cylinder cylindrical stabilizing blades are combined, if the ratio R m / R 0 of the diameter R m of each in-cylinder cylindrical stabilizing blade to the diameter R 0 of the lower cylindrical stabilizing blade is ε m , Formulas 7- (7) to 7- (9) become formulas 7- (10) to 7- (12) as general formulas.
Figure JPOXMLDOC01-appb-M000055
Figure JPOXMLDOC01-appb-M000055
　ここで式７-（１３）と置けば、ＳCは式７-（１４）となり、また式７-（１５）と置けば、式７-（１６）～式７-（１８）を得る。また式６-（２），式６-（３）は、式７-（１９），式７-（２０）となる。 Place here by Formula 7 and (13), S C is if you put the formula 7- (14), and also the formula 7- (15), to obtain the formula 7- (16) to formula 7 (18). Expressions 6- (2) and 6- (3) become Expressions 7- (19) and 7- (20).
Figure JPOXMLDOC01-appb-M000056
Figure JPOXMLDOC01-appb-M000056
　安定翼の形状としては、放射状、円筒状の他、偶数角正多角形筒状およびそれらの組み合わせ等がある。他にも中心軸線方向から見て中心軸線対称な網目状等がある。要するに安定翼の形状としては、プロペラ回転軸に垂直な、どの方向から見てもＦCの大きさが変わらない様な形状であればどんな形状であってもよい。プロペラ回転による反トルクを相殺するには、放射状の安定翼が一番良いと思われる。なぜならば、円筒状あるいは偶数角多角形筒状の安定翼の場合は、実験より、プロペラ回転による反トルクの防止には、殆ど寄与しないことが分かっているからである。また筒状安定翼の形状を円筒ではなく角柱型の角筒にした場合は、上記の偶数角正多角形の部類に入るが、外側の筒状安定翼の場合は、外側にはプロペラ後流が無いので、同じ大きさの筒内筒状安定翼に掛かるＦCの１／２の力のＦCしか働かないことに注意を要する。 As the shape of the stabilizing blade, there are a radial shape, a cylindrical shape, an even angle regular polygonal cylindrical shape, a combination thereof, and the like. In addition, there are a mesh shape that is symmetrical with respect to the central axis when viewed from the central axial direction. The shape of the short stable wing, perpendicular to the propeller shaft, may be in any shape as long as any even when viewed from a direction a shape as does not change the size of the F C. To counteract the counter-torque caused by propeller rotation, a radial stabilizer blade seems to be the best. This is because, in the case of a stable wing having a cylindrical shape or even-numbered polygonal cylindrical shape, it is known from experiments that it hardly contributes to the prevention of the counter torque due to the propeller rotation. In addition, when the cylindrical stabilizer blade is not a cylinder but a prismatic prism, it falls into the above-mentioned even angle regular polygon category, but in the case of the outer cylindrical stabilizer blade, the outer side of the propeller is on the outside. since there is no care must be taken that only work F C 1/2 force F C acting on the cylinder in the cylindrical stabilizing wings of the same size.
　Ｓ８．（筒状安定翼と放射状安定翼を備えた機体のプロペラ回転による反トルクの相殺条件）
　図１１の機体７０は、安定してホバリングを行ったが、プロペラ回転と同方向に正回転した。この正回転を止めるための下部放射安定翼７２の高さのｎ値を求める。 S8. (Anti-torque canceling condition due to propeller rotation of an aircraft equipped with cylindrical stabilizer blades and radial stabilizer blades)
The airframe 70 in FIG. 11 performed hovering stably, but rotated forward in the same direction as the propeller rotation. The n value of the height of the lower radiation stabilizing blade 72 for stopping the forward rotation is obtained.
　図６の様な三角翼の機体６３のプロペラ回転による反トルクの相殺条件（以後、反トルク相殺条件と呼ぶ）は、実施の形態１で示した様に、その機体６３の下部放射状安定翼６１の安定翼の直径単位の枚数をｍとしたとき、その下部放射状安定翼６１の高さのｎ値が式８-（１）で与えられる事である。 The counter-torque canceling condition (hereinafter referred to as the anti-torque canceling condition) of the triangular wing airframe 63 as shown in FIG. 6 is the lower radial stabilizing wing 61 of the airframe 63 as shown in the first embodiment. When the number of stable blades in the diameter unit is m, the n value of the height of the lower radial stabilizer 61 is given by Equation 8- (1).
Figure JPOXMLDOC01-appb-M000057
Figure JPOXMLDOC01-appb-M000057
　しかし図１１の様に下部筒状安定翼７３によりプロペラ後流が下部筒状安定翼７３の外側に漏れなくされた機体７０（放射状安定翼７２の安定翼の直径単位の枚数ｍ＝２）では、下部放射状安定翼７２の高さのｎ値を、式８-（１）に基づきｎ＝πに設定しても、機体７０の回転は止まらずに正回転した。このことは、プロペラ後流による回転モーメント力の方がプロペラ７１が受ける反トルクよりも大きいことを意味する。 However, as shown in FIG. 11, the airframe 70 in which the propeller wake is not leaked to the outside of the lower cylindrical stabilizer wing 73 by the lower cylindrical stabilizer wing 73 (the number m = 2 of the stable blade diameter unit of the radial stabilizer wing 72). Even if the n value of the height of the lower radial stabilizer wing 72 was set to n = π based on the equation 8- (1), the rotation of the airframe 70 did not stop but rotated forward. This means that the rotational moment force caused by the propeller wake is larger than the counter torque that the propeller 71 receives.
　ここで考えられることは、ＦCの定義式（例えば式７-（１６））の中にπが含まれることから、プロペラ後流による回転モーメント力もπ倍になっていると推測できる。その場合の式８-（１）の右辺は１／π倍に修正されるので、機体７０の反トルク相殺条件は、式８-（１）の右辺を１／π倍したものになると推測できる。この推測に基づき、機体７０の下部放射状安定翼７２の高さのｎ値を求めるとｎ＝１となる。このｎ値の下で、重心Ｇおよび外部風圧中心点Ｗ等を調整して、再度、機体７０で実験（上記の第１の実験）を行った。その結果、機体７０は反回転をしながらホバリングを始めた。これは、下部放射状安定翼７２の高さが短すぎたからであると思われる。 What can be considered here is that π is included in the definition formula of F C (for example, Expression 7- (16)), so it can be estimated that the rotational moment force due to the propeller wake is also π times. Since the right side of Equation 8- (1) in that case is corrected to 1 / π times, it can be estimated that the anti-torque canceling condition of the body 70 is 1 / π times the right side of Equation 8- (1). . Based on this estimation, when the n value of the height of the lower radial stabilizer wing 72 of the airframe 70 is obtained, n = 1. Under this n value, the center of gravity G, the external wind pressure center point W, and the like were adjusted, and the experiment (the first experiment described above) was performed again with the body 70. As a result, the fuselage 70 started hovering while rotating counterclockwise. This seems to be because the height of the lower radial stabilizer wing 72 was too short.
　「最新流体工学の基本」（著者：小峯龍男、２００６年４月６日発行）の揚力の項の中に「揚力Ｌ＝ＫρＳＵＶ＝ＫρＳＵ2（α＋β）、ここで比例定数（揚力係数）Ｋは、理論値でπ、実験値で２．７～２．９の値をとります。」という一文があり、揚力に関しては、揚力Ｌがπ倍にならずに２．７～２．９倍位になるということである。 “Lift L = KρSUV = KρSU 2 (α + β), where the proportionality constant (lift coefficient) K” is in the lift section of “Basics of Modern Fluid Engineering” (author: Tatsuo Kominato, issued April 6, 2006) , The theoretical value is π, and the experimental value is 2.7 to 2.9. ”With regard to the lift, the lift L does not become π times but is about 2.7 to 2.9 times. Is to become.
　その一文を参考にすれば、図１１の機体７０で再度実験（上記の第１の実験）をするにあたり、下部放射状安定翼７２の高さのｎ値を式８-（１）の値の１／π倍ではなく１／２．７倍に設定すれば、機体７０の反回転を止めることができると思われる。そしてその実験（第３の実験）の結果、下部放射状安定翼７２の高さのｎ値をπ／２．５４５＝１．２３４に設定すると、機体７０はほとんど反回転せずに安定してホバリングを行った。 Referring to the sentence, in performing the experiment again (the first experiment described above) with the airframe 70 of FIG. 11, the n value of the height of the lower radial stabilizer blade 72 is set to 1 of the value of Equation 8- (1). If it is set to 1 / 2.7 times instead of /.pi. Times, it seems that the anti-rotation of the airframe 70 can be stopped. As a result of the experiment (third experiment), when the n value of the height of the lower radial stabilizer wing 72 is set to π / 2.545 = 1.234, the airframe 70 can be stably hovered with almost no counter-rotation. Went.
　この事実から、前述の三角翼の機体６３の下部放射状安定翼６１の高さのｎ値も約１．２３４位に設定すると、前述の三角翼の機体６３も反回転が止まることを意味する。なぜならば、一般に三角翼の機体の場合、下部放射状安定翼の高さのｎ値が小さいときは、ＦCの力にπ倍効果が出現するからである。 From this fact, if the n value of the height of the lower radial stabilizer wing 61 of the triangular wing airframe 63 is also set to about 1.234, it means that the anti-rotation of the triangular wing airframe 63 also stops. This is because, in general, in the case of a triangular wing airframe, a π-fold effect appears in the force of F C when the n value of the height of the lower radial stabilizer is small.
　Ｓ９．（ＦC＝Ｈπｍｇsinβの修正）
　上述の様にＦCの力にπ倍効果がある場合は、機体の反回転を止めるための計算では、１／π倍ではなく約１／２．５４５倍であった。このことは何を意味するのか。 S9. (Modification of the F C = Hπmgsinβ)
As described above, when there is a π-fold effect on the force of F C , the calculation for stopping the anti-rotation of the aircraft was not about 1 / π but about 1 / 2.545. What does this mean?
　２．５４５／π＝０．８１０１の数字を考えると、２．５４５は、π（従ってπ倍効果）の約８１％になっている。しかし、今までの機体の左右の揺れのモーメントバランスを考えたとき、１００％π倍効果を考慮した計算結果は、実験結果と良く合っていた。ＦCなる力は、機体が動くときの抵抗力であり、機体が垂直姿勢で安定しているときには発生しないこと、および機体が傾き始めたときにも抵抗力として働くことから、プロペラ後流がプロペラから離れた後に機体が動いたとき、プロペラ後流と機体の下部放射状安定翼との間に極わずかな角度（流入角度）βが生じ、その角度分だけのＦCなる力が機体に掛かると考えられる。その角度βが０°のときはＦC＝０であり、その角度βが増大するに連れてＦCも増大する。故に反回転相殺のＦCなる力は、機体を正回転させることができる積極的な力であり、抵抗力であるＦCとは、異種の力である事が分かる。平行流の中の平板の揚力係数と流入角度との関係は、ほぼ直線的な比例関係にあることが知られている。これらの事を総合的に考えたとき、式９-（１）が今までの実験と一番合う反回転相殺ＦCの定義式（近似式）ではないかと思われる。 Considering the number 2.545 / π = 0.8101, 2.545 is about 81% of π (and hence the π-fold effect). However, when considering the moment balance of the left and right shaking of the aircraft until now, the calculation results considering the 100% π times effect matched well with the experimental results. The force of F C is the resistance force when the aircraft moves, it does not occur when the aircraft is stable in a vertical posture, and it works as a resistance force when the aircraft starts to tilt. when the aircraft has moved after leaving the propeller, very small angle (inflow angle) beta is generated between the lower radial stability blades of the propeller slipstream and aircraft, F C becomes the force of only the angle amount is applied to the body it is conceivable that. When the angle β is 0 °, F C = 0, and as the angle β increases, F C also increases. Therefore, it can be seen that the anti-rotation canceling force, F C, is an aggressive force that can rotate the aircraft in the forward direction, and the resistance force, F C , is a different force. It is known that the relationship between the lift coefficient of a flat plate in parallel flow and the inflow angle is in a substantially linear proportional relationship. When these things are considered comprehensively, it seems that Equation 9- (1) is the definition equation (approximate equation) of the counter-rotation cancellation F C that best matches the previous experiments.
Figure JPOXMLDOC01-appb-M000058
Figure JPOXMLDOC01-appb-M000058
　上記のＳ８の実験値から式９-（１）中のβを計算すると、β≒１０．９５°となる。 When β in Equation 9- (1) is calculated from the experimental value of S8, β≈10.95 °.
　このとき、プロペラ後流は下部放射状安定翼に対して１０．９５°の角度で当たっていたことになる。ということは、プロペラの平均ピッチ角度が１０．９５°ということになる。それでは、ピッチ角度の異なるプロペラで同じ実験を行えば式９-（１）を確認できるものと思われる。ここで注意をしなければならないのは、下部放射状安定翼には、実際には常に或る一定の角度でプロペラ後流が当たっているが、下部放射状安定翼の各安定翼の左右の前後で流入角度が逆のため、機体には実際には前後の力は相殺され、ＦCなる力は専ら機体の回転モーメントに関してのみ作用しているということである。 At this time, the propeller wake was hitting the lower radial stabilizer at an angle of 10.95 °. This means that the average pitch angle of the propeller is 10.95 °. Then, it seems that Formula 9- (1) can be confirmed by performing the same experiment with propellers having different pitch angles. It should be noted here that the lower radial stabilizer wing is actually always subjected to the propeller wake at a certain angle, but before and after the left and right sides of each stabilizer wing of the lower radial stabilizer. Since the inflow angle is reversed, the front / rear force is actually canceled out in the airframe, and the force F C acts exclusively with respect to the rotational moment of the airframe.
　β≒１０．９５°なる風を発生させたプロペラのピッチが３インチであったので、次にピッチ５インチのプロペラで同じ実験を行うと、機体の反回転が止まる下部放射状安定翼の高さは、約（π／２．１８）ｒ0であった。このとき、式９-（１）よりβを求めると、β≒１７．８２°となった。 The pitch of the propeller that generated the wind of β ≒ 10.95 ° was 3 inches, so when the same experiment was carried out with a propeller with a pitch of 5 inches, the height of the lower radial stable wing where the anti-rotation of the fuselage stopped Was about (π / 2.18) r 0 . At this time, when β was obtained from Equation 9- (1), β≈17.82 °.
　プロペラのピッチとは、プロペラが１回転して前に進む距離だから、このピッチの異なる２つのプロペラの平均ピッチ角（ねじれ角度）の比が上記の１０．９５°と１７．８２°の比となって現れると思われる。 The pitch of the propeller is a distance that the propeller makes one rotation and moves forward, so the ratio of the average pitch angle (twist angle) of the two propellers with different pitches is the ratio of the above 10.95 ° and 17.82 °. It seems to appear.
　ここで、プロペラの平均ピッチ角度がプロペラのどの位置に現れるのかを考える。プロペラのピッチとは、プロペラが１回転したときにプロペラが進む距離であるので、上記のピッチ３インチの場合とピッチ５インチの場合の実験結果に基づき、式９-（３）および式９-（４）から、それぞれのピッチの場合の平均ピッチ角度の位置を求めることができる（但し、ｒ0＝６インチ＝１５．２４ｃｍ）。 Here, consider the position of the propeller where the average pitch angle of the propeller appears. The pitch of the propeller is the distance traveled by the propeller when the propeller makes one revolution. Therefore, based on the experimental results for the pitch of 3 inches and the pitch of 5 inches, the formulas 9- (3) and 9- From (4), the position of the average pitch angle for each pitch can be obtained (where r 0 = 6 inches = 15.24 cm).
Figure JPOXMLDOC01-appb-M000059
Figure JPOXMLDOC01-appb-M000059
　この結果から、プロペラの中心軸から半径の８２．５％位の所のピッチ角度がそのプロペラの実効角度であると言える。上記のピッチ５インチのプロペラの最先端のピッチ角度を求めると、式９-（５）となる。 From this result, it can be said that the pitch angle at a position of about 82.5% of the radius from the central axis of the propeller is the effective angle of the propeller. When the most advanced pitch angle of the 5-inch pitch propeller is obtained, Equation 9- (5) is obtained.
Figure JPOXMLDOC01-appb-M000060
Figure JPOXMLDOC01-appb-M000060
　この結果から、揚力が最大となる角度１５°に殆ど近い事が分かる。そしてこのプロペラ使用時の反トルクを相殺する下部放射状安定翼の高さのｎ値は、式９-（６）であった。 From this result, it can be seen that it is almost close to the angle of 15 ° where the lift is maximum. The n-value of the height of the lower radial stabilizer that cancels out the counter-torque when using the propeller was expressed by Equation 9- (6).
Figure JPOXMLDOC01-appb-M000061
Figure JPOXMLDOC01-appb-M000061
　一般にプロペラの実効角度（流入角度）をβTとすれば、π倍効果が現れている状態でのプロペラ回転による反トルクを相殺する下部放射状安定翼の高さのｎ値は（そのときのｎ値をｎTとして置く）、式９-（２）となる。 In general, if the effective angle (inflow angle) of the propeller is β T , the n value of the height of the lower radial stabilizer blade that cancels the counter-torque due to the rotation of the propeller in the state where the π-fold effect appears (the n value at that time) The value is set as n T ), resulting in Equation 9- (2).
Figure JPOXMLDOC01-appb-M000062
Figure JPOXMLDOC01-appb-M000062
　尚、下部放射状安定翼の直径単位の枚数がｍ枚の場合は、式９-（９）となる。 In addition, when the number of diameter units of the lower radial stabilizer is m, Expression 9- (9) is obtained.
Figure JPOXMLDOC01-appb-M000063
Figure JPOXMLDOC01-appb-M000063
　尚、ピッチ５インチのプロペラの場合は、反回転を止めるための下部放射状安定翼のｎ値には幅があり、約１．４～１．６位の間のｎ値のときに、機体は、１回のホバリング中に正回転と反回転とを繰り返していた。上記のｎ値１．４４という値は、一番安定していると思われるｎ値の値である。このプロペラの先端のピッチ角度がほぼ１５°になっていることにより、先端から内側のピッチ角度は１５°を超えていることになる。一般の揚力理論では、仰角１５°を越えると、揚力は減少して行くので、このプロペラの場合、プロペラの内側に行くほど、ピッチ角度が１５°よりも増加して行き、それとは逆に、揚力は減少して行くことになる。この事がプロペラの反トルクを相殺するための下部放射状安定翼のｎ値に幅を持たせているものと思われる。 In the case of a propeller with a pitch of 5 inches, the n-value of the lower radial stabilizer for stopping counter-rotation has a width, and when the n-value is between about 1.4 to 1.6, During one hovering, the forward rotation and the reverse rotation were repeated. The value of n value 1.44 is the value of n value that seems to be the most stable. Since the pitch angle at the tip of this propeller is approximately 15 °, the pitch angle on the inner side from the tip exceeds 15 °. In general lift theory, the lift decreases when the elevation angle exceeds 15 °. In the case of this propeller, the pitch angle increases more than 15 ° toward the inside of the propeller, and conversely, Lift will go down. This seems to give a wide range to the n value of the lower radial stabilizer for offsetting the propeller's counter torque.
　Ｓ１２．（ＦC＝ｎπｍｇsinβの物理的意味）
　上述の様にＦCの式は全てプロペラの直径ｒ0が計算の基本となっている。プロペラの回転中の一瞬一瞬でプロペラの直径ｒ0全体が機体の重量ｍｇを支えていることを考えれば、πｍｇの意味は、プロペラ１回転のプロペラの浮力の総和である。 S12. (Physical meaning of F C = nπmgsinβ)
As described above, all the formulas of F C are based on the propeller diameter r 0 . Considering that the entire propeller diameter r 0 supports the weight mg of the aircraft in a moment during the rotation of the propeller, the meaning of π mg is the sum of the buoyancy of the propeller with one rotation of the propeller.
　ｎはプロペラ直径ｒ0の倍数係数なので、プロペラ後流の拡がり角度がどうであれ、安定翼がプロペラ後流全てを受け留められるものである限り、安定翼に掛かるプロペラ後流による風圧力（揚力）ＦCは、安定翼の高さｎｒ0に正比例して増大する。 Since n is a multiple coefficient of the propeller diameter r 0 , wind pressure (lift force) generated by the propeller wake applied to the stable wing as long as the stable wing can receive all the wake behind the propeller regardless of the spread angle of the propeller wake. ) F C increases in direct proportion to the height nr 0 of the stabilizer blade.
　しかもプロペラ１回転の力の総和πｍｇがｎ値に比例して安定翼に掛かることになる。このことは、プロペラ１回転の円の内にある全総力（πｍｇ）がほとんど一瞬にプロペラから風となって離れ、下に間断無く流れてゆき常に安定翼に対して力を及ぼしていること、即ち換言すれば、口径ｒ0の穴から噴出する間断無い風が安定翼に対して口径ｒ0のｎ倍に比例した力を与えていることと同じである。 Moreover, the total force πmg of one propeller rotation is applied to the stable blade in proportion to the n value. This means that the total force (πmg) within the circle of one rotation of the propeller is almost instantaneously separated from the propeller by the wind, and flows downward without interruption. or in other words, is the same as that interruption no wind ejected from holes of diameter r 0 is given a force proportional to n times the diameter r 0 on the stability wing.
　風の発生がプロペラに依らない風（例えば爆発による風）が、口径ｒ0の穴から噴出しているとき、その穴全体から噴出する風の一瞬の風圧力をＰeとすれば、筒状安定翼内の安定翼に掛かるプロペラ後流による風圧力ＦCの基本式である式７-（１６）は、式１２-（２）と書き換える事ができると思われる。 Wind generation of wind does not depend on the propeller (e.g. wind by explosion) is, when you are ejected from the hole of diameter r 0, if the wind pressure of the moment of wind ejected from the entire hole and P e, tubular Formula 7- (16), which is the basic formula of the wind pressure F C caused by the propeller wake on the stabilizer blade in the stabilizer blade, can be rewritten as Equation 12- (2).
Figure JPOXMLDOC01-appb-M000064
Figure JPOXMLDOC01-appb-M000064
　そしてＦWおよびＷの式も式１２-（３）～式１２-（５）の様になる。 The formulas for FW and W are also given by formulas 12- (3) to 12- (5).
Figure JPOXMLDOC01-appb-M000065
Figure JPOXMLDOC01-appb-M000065
　式１２-（２）～式１２-（５）は全てπ倍効果が出現している場合の式である。π倍効果が出現していないときのＦC、ＦWおよびＷは、式１２-（６）～式１２-（９）の様になる。 Expressions 12- (2) to 12- (5) are expressions when the π-fold effect appears. F C , F W, and W when the π-fold effect does not appear are expressed by Equations 12- (6) to 12- (9).
Figure JPOXMLDOC01-appb-M000066
Figure JPOXMLDOC01-appb-M000066
　以上からこれまでの議論は、プロペラ機だけではなく口径ｒ0の穴から噴出される風によるホバリング機（例えばロケット、ジェット機およびガス噴射等によるホバリング機）にも応用可能である事が分かる。 From the above, it can be seen that the discussion so far can be applied not only to propeller aircraft but also to hovering devices using wind blown from a hole having a diameter r 0 (for example, rockets, jet aircraft, and hovering devices using gas injection, etc.).
　Ｓ１３．（下部筒状安定翼をプロペラよりも上に延長した場合）
　図１２の機体８０は、プロペラ８１と、プロペラ８１の下側に配置された側面視矩形状で例えば十字状の下部放射状安定翼８２と、下部放射状安定翼８２の周囲を囲繞する様に同軸線状に配設され、その下端が下部放射状安定翼８２と同じ高さで且つその上端がプロペラ８１よりも上側に延長された筒状安定翼８３と、下部放射状安定翼８２に配設されたプロペラ用の駆動部（不図示）とを備えている。 S13. (When the lower cylindrical stabilizer is extended above the propeller)
The airframe 80 in FIG. 12 has a propeller 81, a rectangular shape in a side view disposed below the propeller 81, for example, a cross-shaped lower radial stabilizer wing 82, and a coaxial line so as to surround the periphery of the lower radial stabilizer wing 82 A cylindrical stabilizer wing 83 whose lower end is the same height as the lower radial stabilizer wing 82 and whose upper end is extended above the propeller 81, and a propeller disposed on the lower radial stabilizer wing 82. Drive unit (not shown).
　この機体８０の様に、筒状安定翼８３の高さをプロペラ８１の上側まで延長した場合、筒状安定翼８３のプロペラ後流による風圧中心点Ｈは、どのような位置に出現し、その位置に作用するＦCなる力（揚力）はどういう式で表されるのかを実験で確認した。 When the height of the cylindrical stabilizer wing 83 is extended to the upper side of the propeller 81 as in the airframe 80, the wind pressure center point H due to the wake of the propeller of the cylindrical stabilizer wing 83 appears at any position. F C becomes the force acting on the position (lift) was confirmed experimentally whether represented by what formula.
　この実験を行うに当たり、予想として、筒状安定翼８３のプロペラ後流による風圧中心点Ｈは、筒状安定翼８３の上端から距離ｈ0ｒ0／４下がった所に位置し、そしてそこに作用するＦCなる力（揚力）は式１３-（１）で表されると仮定した。式９-（１）を採用しない理由は、抵抗力の計算時には、（１－sinβ）の要素が無い式１３-（１）の様な式が、経験上、実験値と良く合うからである。 In making this experiment, as expected, the wind pressure center point H by the propeller slipstream of the cylindrical stabilizing wing 83 is located from the upper end of the cylindrical stabilizing wing 83 a distance h 0 r 0/4 lowered place, and there F C becomes the force acting (lift) was assumed to be represented by the formula 13- (1). The reason why Expression 9- (1) is not adopted is that, when calculating the resistance force, an expression like Expression 13- (1) having no element of (1-sinβ) matches experimental values well from experience. .
Figure JPOXMLDOC01-appb-M000067
Figure JPOXMLDOC01-appb-M000067
　機体８０の重心Ｇを、筒状安定翼８３におけるプロペラ後流による風圧中心点Ｈと下部放射状安定翼８２におけるプロペラ後流による風圧中心点Ｃとの合力の風圧中心点（総合風圧中心点）Ｃ0から式７-（２０）で表される距離だけ離した点に置き、点Ｗの位置は、式７-（１８）で与えられる距離よりも少し大きく取り、且つｎ≒１，４４として機体８０をホバリングさせた。尚、上記の総合風圧中心点Ｃ0は、各点Ｈ，Ｃ間の距離を式１３-（２）の比率で割った点である。 The center of gravity G of the fuselage 80 is set to a wind pressure center point (total wind pressure center point) C of the resultant force between the wind pressure center point H caused by the propeller wake in the cylindrical stabilizer wing 83 and the wind pressure center point C caused by the propeller wake in the lower radial stabilizer wing 82. Place it at a point separated from 0 by the distance represented by Expression 7- (20), and the position of the point W is slightly larger than the distance given by Expression 7- (18), and n≈1,44. 80 was hovered. The total wind pressure central point C 0 is a point obtained by dividing the distance between the points H and C by the ratio of the equation 13- (2).
Figure JPOXMLDOC01-appb-M000068
Figure JPOXMLDOC01-appb-M000068
　尚、ｎは、下部放射状安定翼８２の高さのｎ値である。またここでは、下部放射状安定翼８２におけるプロペラ後流による風圧中心点Ｃは、揚力の一般理論に従って、下部放射状安定翼８２の上端から距離ｎｒ0／４下がった位置になる場合を考えている。 Note that n is the n value of the height of the lower radial stabilizer wing 82. Here also, the wind pressure center point C by the propeller slipstream at lower radial stabilizing wings 82, according to the general theory of lift, and consider the case comprising an upper end a distance nr 0/4 down position of the lower radial stabilization wings 82.
　その結果は予想通り、機体８０は、非常に安定してホバリングし、反回転も生じなかった。今回の実験も無風の中で行った。今回の実験では外部風圧中心点Ｗが式７-（９）で表される点よりも少し下にあったが、無風の中の実験では、今までで一番安定したホバリングを行った様に見えた。 As expected, the aircraft 80 hovered very stably and no anti-rotation occurred. This experiment was also conducted in the absence of wind. In this experiment, the external wind pressure center point W was slightly below the point expressed by Equation 7- (9). However, in the experiment without wind, it seems that the most stable hovering has been performed so far. Looked.
　この機体８０を少し風のある中でホバリングさせると、機体８０はほとんど平行に風に流され始め、風が止まるとまたその移動先で安定したホバリングを行っていた。この機体８０では各点Ｇ，Ｗの平衡点が少しずれていたので、外風に流されて平行移動しているときには少し傾き加減ではあったが計算通りの安定性を見せてくれた。 When the airframe 80 was hovered in a little wind, the airframe 80 started to be blown almost in parallel with the wind, and when the wind stopped, it was performing stable hovering at the destination. In this airframe 80, the equilibrium points of the points G and W were slightly deviated from each other, so when they were moved by the outside wind and moved in parallel, the slope was slightly adjusted but the stability as calculated.
　この実験結果から途切れのない筒状安定翼８３の中にプロペラ８１を配置したとき、プロペラ８１の上側と下側においては、筒状安定翼８３の断面を単位時間当たり通過する空気の流量が等しいことが分かる。その結果、筒状安定翼８３におけるプロペラ後流による風圧中心点Ｈも予想通り、筒状安定翼８３の上端から距離ｈ0ｒ0／４下がった位置に存在し、重心Ｇ周りのモーメントが平衡する点から式７-（１９）または式７-（２０）で表される点に重心Ｇを配置すれば、非常に安定したホバリングを得られることが分かった。 From this experimental result, when the propeller 81 is disposed in the cylindrical stabilizer blade 83 without interruption, the flow rate of air passing through the section of the cylindrical stabilizer blade 83 per unit time is equal on the upper side and the lower side of the propeller 81. I understand that. As a result, expected also wind pressure center point H by the propeller slipstream of the tubular stabilizer blades 83, there from the upper end of the cylindrical stabilizing wing 83 a distance h 0 r 0/4 down position, moment about the center of gravity G is balanced From this point, it was found that a very stable hovering can be obtained if the center of gravity G is arranged at a point represented by Expression 7- (19) or Expression 7- (20).
　Ｓ１４．（色々な種類の安定翼の組み合わせによる姿勢制御装置）
　図１３の機体（姿勢制御装置）９０は、筒状安定翼９１と、筒状安定翼９１の内部に筒状安定翼９１の中心軸線９２上に沿って同軸線状に配設された１つ以上（ここでは２つ）の中心軸線対称で例えば側面視矩形の放射状安定翼９３，９４と、筒状安定翼９１の内部に同軸線状に配設された１つ以上（ここでは２つ）の筒内筒状安定翼９５，９６と、筒状安定翼（例えば９１）の中心軸線上に配設されたプロペラ９７と、放射状安定翼（例えば９３）に配設されたプロペラ用の駆動部（不図示）とを備える。ここでは、各筒状安定翼９１，９５，９６は例えば円筒状に形成されている。またプロペラ９７の直径は、筒状安定翼９１の口径よりも小さいものとする。 S14. (Attitude control device by combining various types of stabilizing blades)
The fuselage (attitude control device) 90 of FIG. 13 is a cylindrical stabilizer wing 91 and one coaxially arranged along the central axis 92 of the cylindrical stabilizer wing 91 inside the cylindrical stabilizer wing 91. The radial stabilizers 93 and 94 that are symmetrical with respect to the central axis as described above (two in this example), for example, are rectangular in side view, and one or more (two in this case) that are coaxially disposed inside the cylindrical stabilizer 91. In-cylinder cylindrical stabilizer blades 95, 96, a propeller 97 disposed on the central axis of the cylindrical stabilizer blade (eg 91), and a propeller drive unit disposed on the radial stabilizer blade (eg 93) (Not shown). Here, each of the cylindrical stabilizing blades 91, 95, 96 is formed in a cylindrical shape, for example. The diameter of the propeller 97 is assumed to be smaller than the diameter of the cylindrical stabilizing blade 91.
　放射状安定翼９３は、例えば、２枚の安定翼を有しており、筒状安定翼９１内の上段に配設されている。また放射状安定翼９４は、例えば、４枚の安定翼を有しており、筒状安定翼９１内の下段に配設されている。各筒内筒状安定翼９５，９６は、例えば、それぞれ異なる口径を有しており、下段放射状安定翼９４と交差する様に筒状安定翼９１内の下段に配設されている。 The radial stabilizer wing 93 has, for example, two stabilizer wings, and is arranged in the upper stage in the cylindrical stabilizer wing 91. The radial stabilizer wing 94 has, for example, four stabilizer blades, and is disposed in the lower stage in the cylindrical stabilizer blade 91. For example, the in-cylinder cylindrical stabilizing blades 95 and 96 have different diameters, and are disposed in the lower stage in the cylindrical stabilizing blade 91 so as to intersect the lower radial stabilizing blade 94.
　この機体９０では、筒状安定翼９１と上段放射状安定翼９３との各々におけるプロペラ後流による風圧中心点Ｈ0，Ｃ1を互いに一致させ、また下段放射状安定翼９４と各筒内筒状安定翼９５，９６との各々におけるプロペラ後流による風圧中心点Ｃ2，Ｈ1，Ｈ2を互いに一致させ、上記の各風圧中心点Ｈ0，Ｃ1の一致点に掛かるプロペラ後流の風圧力の合力ＦC1の大きさと上記の各風圧中心点Ｃ2，Ｈ1，Ｈ2の一致点に掛かるプロペラ後流の風圧力の合力ＦC2の大きさとの比率を調整して、プロペラ後流総合風圧中心点Ｃ0から機体９０の重心までの距離ｎGC0と、外部風圧中心点Ｗから機体９０の重心までの距離ｎGWと、プロペラ回転軸の固定点０から機体９０の重心までの距離ｎGとの３要素が式１９-（５）を満たす様に調整されている。 In this airframe 90, the wind pressure center points H 0 and C 1 caused by the propeller wakes in the cylindrical stabilizer wing 91 and the upper radial stabilizer wing 93 are made to coincide with each other, and the lower radial stabilizer wing 94 and each cylindrical stabilizer Wind pressure center points C 2 , H 1 , H 2 due to the propeller wake at each of the blades 95, 96 are made to coincide with each other, and the wind pressure at the wake of the propeller applied to the coincidence of the wind pressure center points H 0 , C 1 described above. of the resultant force F magnitude and the wind pressure center point C 2 of the above C1, by adjusting the ratio between the size of the H 1, the resultant force of the wind pressure of H 2 applied to the matching point propeller slipstream F C2, after propeller stream Overall The distance n GC0 from the wind pressure center point C 0 to the center of gravity of the fuselage 90, the distance n GW from the external wind pressure center point W to the center of gravity of the fuselage 90, and the distance n from the fixed point 0 of the propeller rotating shaft to the center of gravity of the fuselage 90 3 elements with the G is adjusted so as to satisfy the formula 19 (5) That.
　ＦC1とＦC2との大きさの比率を調整する方法は、例えば下記（１）～（８）が考えられる。 For example, the following (1) to (8) can be considered as a method of adjusting the ratio of the sizes of F C1 and F C2 .
　（１）放射状安定翼９３，９４の安定翼の枚数を増減する
　（２）放射状安定翼９３，９４の高さを増減する
　（３）筒状安定翼９１の高さを増減する
　（４）筒内筒状安定翼９５，９６の高さを増減する
　（５）筒内筒状安定翼９５，９６の直径を増減する
　（６）筒内筒状安定翼の個数を増減する
　（７）各安定翼９３，９４，９５，９６の位置を調整する
　（８）幾つかの安定翼（放射状安定翼および／または筒内筒状安定翼）を追加する。 (1) Increase or decrease the number of stabilizing blades of the radial stabilizing blades 93, 94. (2) Increase or decrease the height of the radial stabilizing blades 93, 94. (3) Increase or decrease the height of the cylindrical stabilizing blade 91. Increase / decrease the height of the inner cylindrical stabilizer blades 95, 96 (5) Increase / decrease the diameter of the cylindrical stabilizer blades 95, 96 (6) Increase / decrease the number of in-cylinder stabilizer blades (7) Each stability Adjust the position of the blades 93, 94, 95, 96. (8) Add some stabilizer blades (radial stabilizer blades and / or in-cylinder cylindrical stabilizer blades).
　その他、筒状安定翼９１の外側に更に通常の翼（即ち板状の翼）を追加して外部風圧中心点Ｗの位置を調整することもできる。 In addition, the position of the external wind pressure center point W can be adjusted by adding a normal blade (that is, a plate-shaped blade) to the outside of the cylindrical stabilizing blade 91.
　この様に構成した機体９０は、全く安定してホバリングを行い、外部風による影響に対しても抵抗力を持ち、驚くべき安定性を保ちつつホバリングを行うことができる。 The airframe 90 configured in this manner performs hovering with stability, has resistance to the influence of external wind, and can perform hovering while maintaining surprising stability.
　ここで機体９０のプロペラ回転による反回転を止めるためには、各放射状安定翼９３，９４のｎ値ｎ1，ｎ2および各放射状安定翼９３，９４の安定翼の直径単位の枚数ｍ1，ｍ2を、式１４-（１）を満たす様に調整すれば良い。尚、βTは、プロペラ後流が放射状安定翼９３，９４の各安定翼の主面に当たる流入角度（換言すればプロペラ９７の実効角度）である。尚、一般にはｉ番目の放射状安定翼のｎ値をｎiとし、安定翼の枚数をｍiとすれば良い。 Here, in order to stop the anti-rotation due to the propeller rotation of the airframe 90, the n value n 1 , n 2 of each radial stabilizer wing 93, 94 and the number m 1 of the diameter units of the stable blade of each radial stabilizer wing 93, 94, m 2 may be adjusted to satisfy Expression 14- (1). Note that β T is an inflow angle (in other words, an effective angle of the propeller 97) at which the wake of the propeller hits the main surface of each of the stabilizing blades of the radial stabilizing blades 93 and 94. In general the n-value of the i-th radial stability blades at n i, may be the number of stable wing and m i.
Figure JPOXMLDOC01-appb-M000069
Figure JPOXMLDOC01-appb-M000069
　以上の様に構成された機体（姿勢制御装置）９０によれば、筒状安定翼９１により、放射状安定翼９３，９４への風流が筒状安定翼９１の外側に拡がる事を防止できると共にその風流を一様にできるので、風流の中において当該機体９０の安定性を向上できる。 According to the fuselage (attitude control device) 90 configured as described above, the cylindrical stabilizer wing 91 can prevent the wind flow to the radial stabilizer wings 93 and 94 from spreading outside the cylindrical stabilizer wing 91. Since the airflow can be made uniform, the stability of the aircraft 90 can be improved in the airflow.
　また筒状安定翼９１の内部に同軸線状に１つ以上の筒内筒状安定翼９５，９６を更に備えるので、更に風流の中において機体９０の安定性を向上できる。 Further, since one or more in-cylinder cylindrical stabilizing blades 95 and 96 are coaxially provided inside the cylindrical stabilizing blade 91, the stability of the airframe 90 can be further improved in the wind flow.
　またプロペラ９７を備えるので機体９０を推進装置として利用でき、安定した飛行を行う事ができる。 Also, since the propeller 97 is provided, the airframe 90 can be used as a propulsion device, and stable flight can be performed.
　また放射状安定翼９３，９４および筒内筒状安定翼９５，９６は、機体９０の重心Ｇ、総合風圧中心点Ｃ0、外部風圧中心点Ｗおよびプロペラ回転軸の固定点０が互いに式１９-（５）を満たす様に配置されるので、機体９０の横揺れを防止できる。 The radial stabilizer blades 93 and 94 and the in-cylinder cylindrical stabilizer blades 95 and 96 have the center of gravity G of the fuselage 90, the total wind pressure center point C 0 , the external wind pressure center point W, and the fixed point 0 of the propeller rotation shaft, respectively. Since it arrange | positions so that (5) may be satisfy | filled, the rolling of the body 90 can be prevented.
　また式１４-（１）を満たす様に設計されるので、プロペラ回転による反トルクを相殺でき、機体９０がプロペラ回転により反回転する事を防止できる。 Also, since it is designed to satisfy the formula 14- (1), the counter-torque caused by the propeller rotation can be offset and the aircraft 90 can be prevented from counter-rotating due to the propeller rotation.
　Ｓ１５．（姿勢制御装置の使用例）
　図１４の機体（姿勢制御装置）１１０は、上記Ｓ１４の機体９０を２つ（以後、それらを機体９０ａ，９０ｂと呼ぶ）組み合わせたものである。より詳細には、この機体１１０は、それぞれその吸気側開口端を上側に向けると共にその排気側開口端を下側に向け、且つ互いの吸気側開口端を互いの対向方向に傾斜（傾斜角度０°）させる様にして、互いに間隔空けて配置された上記の２つの機体９０ａ，９０ｂと、各機体９０ａ，９０ｂを相互連結する連結部材１１１とを備える。 S15. (Usage control device usage example)
The airframe (attitude control device) 110 in FIG. 14 is a combination of two airframes 90 of S14 (hereinafter referred to as airframes 90a and 90b). More specifically, the airframe 110 has its intake-side open end directed upward, its exhaust-side open end directed downward, and its intake-side open ends inclined in opposite directions (inclination angle 0). The above-mentioned two airframes 90a and 90b arranged at a distance from each other and a connecting member 111 for interconnecting the airframes 90a and 90b.
　連結部材１１１は、その中心点Ｇ0で屈曲された略Ｖ字状の例えば棒状に形成されており、その一端に機体９０ａが配設され、その他端に機体９０ｂが配設されている。より詳細には、連結部材１１１の一端の延長線が機体９０ａの中心軸線に直交し且つ機体９０ａの重心Ｇ1を通過する様に、連結部材１１１の一端に機体９０ａが配設されている。同様に、連結部材１１１の他端の延長線が機体９０ｂの中心軸線に直交し且つ機体９０ｂの重心Ｇ2を通過する様に、連結部材１１１の他端に機体９０ｂが配設されている。 The connecting member 111 is formed in a substantially V-shape, for example, a rod shape bent at the center point G 0 , and a body 90 a is disposed at one end and a body 90 b is disposed at the other end. More specifically, as an extension of one end of the connecting member 111 passes through the center of gravity G 1 of and body 90a orthogonal to the central axis of the body 90a, the aircraft 90a to one end of the connecting member 111 is disposed. Similarly, as an extension of the other end of the connecting member 111 passes through the center of gravity G 2 orthogonally and body 90b to the center axis of the body 90b, aircraft 90b to the other end of the connecting member 111 is disposed.
　尚、各機体９０ａ，９０ｂのプロペラ９７の回転方向を逆向きに設定する場合は、各機体９０ａ，９０ｂの各安定翼（円筒状安定翼および放射状安定翼）を、その機体９０ａ，９０ｂの重量の許される範囲内で、高さ、安定翼の枚数および安定翼の個数等を大きくして、機体１１０の横揺れの抵抗力を最大限にしておくことが望ましい。 When the rotation direction of the propeller 97 of each of the airframes 90a and 90b is set in the reverse direction, the weight of each of the airframes 90a and 90b is the weight of the airframe 90a and 90b. It is desirable to maximize the rolling resistance of the fuselage 110 by increasing the height, the number of stabilizing blades, the number of stabilizing blades, and the like within the allowable range.
　尚、図１４中の点Ｐ1，Ｐ2はそれぞれ、機体９０ａ，９０ｂのプロペラ９７の中心点であり、αは、線分Ｇ1Ｇ0（および線分Ｇ2Ｇ0）と水平方向との間の角度であり、βは、線分Ｐ1Ｇ0（および線分Ｐ2Ｇ0）と鉛直方向との間の角度であり、θは、線分Ｇ1Ｇ0と線分Ｐ1Ｇ0との間の角度および線分Ｇ2Ｇ0と線分Ｐ2Ｇ0との間の角度である。 Note that points P 1 and P 2 in FIG. 14 are the center points of the propellers 97 of the airframes 90a and 90b, respectively, and α is the line segment G 1 G 0 (and line segment G 2 G 0 ) and the horizontal direction. Is the angle between the line segment P 1 G 0 (and the line segment P 2 G 0 ) and the vertical direction, and θ is the line segment G 1 G 0 and the line segment P 1. it is the angle between the angle and line G 2 G 0 and the line segment P 2 G 0 between G 0.
　ここで、線分Ｇ0Ｇ1＝線分Ｇ0Ｇ2＝Ｌ、線分Ｇ0Ｐ1＝線分Ｇ0Ｐ2＝ｌとし、各機体９０ａ，９０ｂの推進力をそれぞれＦ，Ｆ’（但し、Ｆ＝Ｆ’）とし、Ｆ，Ｆ’の重心Ｇ0周りのモーメントをそれぞれＭ，Ｍ’とすると、式１５-（１）となる。 Here, line segment G 0 G 1 = line segment G 0 G 2 = L, line segment G 0 P 1 = line segment G 0 P 2 = l, and the propulsive forces of the airframes 90a and 90b are F and F ', respectively. (Where F = F ′), and the moments around the center of gravity G 0 of F and F ′ are M and M ′, respectively, Equation 15- (1) is obtained.
Figure JPOXMLDOC01-appb-M000070
Figure JPOXMLDOC01-appb-M000070
　この機体１１０が角度γだけ一方の機体（例えば９０ａ）側に傾いた状態を考えると、Ｍ，Ｍ’は、式１５-（２）となる。 Considering a state in which the airframe 110 is inclined toward one airframe (for example, 90a) by an angle γ, M and M ′ are expressed by Expression 15- (2).
Figure JPOXMLDOC01-appb-M000071
Figure JPOXMLDOC01-appb-M000071
　従って、機体１１０を元に戻そうとするモーメント〔Ｍ〕を求めると、式１５-（３）となる。 Therefore, when the moment [M] for returning the airframe 110 to the original position is obtained, Expression 15- (3) is obtained.
Figure JPOXMLDOC01-appb-M000072
Figure JPOXMLDOC01-appb-M000072
　また機体１１０が角度γだけ傾いたときのＦとＦ’の合力〔Ｆ〕は、式１５-（４）となる。 Also, the resultant force [F] of F and F ′ when the airframe 110 is tilted by the angle γ is expressed by Equation 15- (4).
Figure JPOXMLDOC01-appb-M000073
Figure JPOXMLDOC01-appb-M000073
　式１５-（３）および式１５-（４）より式１５-（５）を得る。 Equation 15- (5) is obtained from Equation 15- (3) and Equation 15- (4).
Figure JPOXMLDOC01-appb-M000074
Figure JPOXMLDOC01-appb-M000074
　式１５-（５）の右辺のｌsinβtanαは線分Ｇ0Ｏの長さなので、点Ｇ0から上に線分Ｇ0Ｏ上がった点ＯにＦとＦ’の合力〔Ｆ〕の作用点があることになる。そして機体１１０が角度γだけ傾くと、〔Ｆ〕・線分Ｇ0Ｏ・sinγで与えられる偶力（復元力）が発生し、その偶力により機体１１０の傾きが元の鉛直方向に復元される。 Since lsinβtanα of the right side of equation 15- (5) is a length of the line segment G 0 O, the point of application of resultant force F and F 'in O line G 0 O rose point on from point G 0 (F) is There will be. When the body 110 is tilted by an angle γ, a couple (restoring force) given by [F] · line segment G 0 O · sin γ is generated, and the tilt of the body 110 is restored to the original vertical direction by the couple. The
　尚、図１４でｌ＝Ｌ／cosθなので、式１５-（５）は式１５-（７）の様に表わせる。 In FIG. 14, since l = L / cos θ, Expression 15- (5) can be expressed as Expression 15- (7).
Figure JPOXMLDOC01-appb-M000075
Figure JPOXMLDOC01-appb-M000075
　式１５-（５）および式１５-（７）より復元力を大きくする要素として、下記（ｉ）（ii）が考えられる。
（ｉ）ｌまたはＬを大きくする
（ii）α，β，θを共に大きくする（但し、α＋β＋θ＝π／２）（尚、αを大きくすると、推進力が減少するので注意を要する）
　以上の様な機体（姿勢制御装置）１１０であれば、横揺れに対して更に安定したホバリングを得ることができる。問題は、機体９０ａ，９０ｂの推進力Ｆ，Ｆ’に差があり過ぎて、その差が復元力を越えてしまう場合であるが、この問題は、事前に機体９０ａ，９０ｂの動力を調整しておけば良いことである。 The following (i) and (ii) are conceivable as factors that increase the restoring force from the equations 15- (5) and 15- (7).
(I) Increase l or L (ii) Increase both α, β, and θ (however, α + β + θ = π / 2) (Note that increasing α reduces propulsive force and requires caution)
With the airframe (posture control device) 110 as described above, it is possible to obtain more stable hovering against rolling. The problem is when the propulsive forces F and F ′ of the airframes 90a and 90b are too different and the difference exceeds the restoring force. This problem is caused by adjusting the power of the airframes 90a and 90b in advance. It is a good thing.
　以上の説明では、この機体１１０の各機体９０ａ，９０ｂの対向方向の揺れに対する復元力に関して説明したが、この機体１１０の各機体９０ａ，９０ｂの対向方向に直交する方向の揺れに対する復元力に関しては、図１４の点Ｐ（線分Ｐ1Ｐ2の中点）に各機体９０ａ，９０ｂの推進力Ｆ，Ｆ’の合力〔Ｆ〕（＝２Ｆcosα）が作用するので、その場合の揺れ（即ち各機体９０ａ，９０ｂの対向方向に直交する方向の揺れ）に対する復元力の方が、各機体９０ａ，９０ｂの対向方向の揺れに対する復元力よりも格段に大きくなることが分かる。 In the above description, the restoring force with respect to the shaking of the airframes 90a and 90b in the opposing direction of the airframe 110 has been described, but the restoring force with respect to the shaking of the airframe 110 in the direction perpendicular to the opposing directions of the airframes 90a and 90b. 14, the resultant force [F] (= 2Fcosα) of the propulsive forces F and F ′ of the airframes 90a and 90b acts on the point P (the midpoint of the line segment P 1 P 2 ). It can be seen that the restoring force with respect to the shaking in the direction orthogonal to the opposing direction of each aircraft 90a, 90b is significantly greater than the restoring force with respect to the shaking in the opposing direction of each aircraft 90a, 90b.
　よって機体１１０を２個使用し、それらを互いの点Ｇ0で平面視で直交する様に組み合わせた機体（姿勢制御装置）を考えれば、その機体は、機器１１０を１個だけで使用する場合よりも復元力が格段に増大することは明らかである。故にこの機器１１０は２個以上組み合わせて使用されることが望ましい。 Therefore, when considering an airframe (posture control device) that uses two airframes 110 and combines them so that they are orthogonal to each other at a point G 0 , the airframe uses only one device 110. It is clear that the resilience increases significantly. Therefore, it is desirable to use two or more devices 110 in combination.
　Ｓ１７．（図１３のプロペラが無い場合の姿勢制御装置）
　図１３において、プロペラ９３（および駆動部）を除いた場合の機体（姿勢制御装置）９０ｃを考える。 S17. (Attitude control device without propeller in FIG. 13)
In FIG. 13, consider an airframe (posture control device) 90 c without the propeller 93 (and the drive unit).
　例えば、この機体９０ｃを一般的な飛行機の前後部および両翼に飛行機の進行方向に向けて取り付けておけば、飛行機の上下左右の安定性が格段に増大する。即ち飛行機が飛んでいるとき、この機体９０ｃに前方（機体９０ｃの中心軸方向）から高速度で風が進入し、その状況で、この機体９０ｃが横方向（その径方向）に揺れると、その前方からの風流が機体９０ｃの筒状安定翼９１および安定翼９３，９４，９５，９６の横方向の揺れに対する抵抗となり、機体９０ｃ（従って当該飛行機）の安定性が飛躍的に向上する。 For example, if this airframe 90c is attached to the front and rear parts and both wings of a general airplane in the direction of travel of the airplane, the stability of the airplane in the vertical and horizontal directions will be greatly increased. That is, when an airplane is flying, wind enters the aircraft 90c from the front (in the direction of the central axis of the aircraft 90c) at a high speed, and in this situation, if the aircraft 90c sways in the lateral direction (its radial direction) The wind flow from the front serves as a resistance against lateral shaking of the cylindrical stabilizer wing 91 and the stabilizer wings 93, 94, 95, 96 of the fuselage 90c, and the stability of the fuselage 90c (and thus the airplane) is greatly improved.
　機体９０ｃの放射状安定翼９３，９４の枚数および筒状安定翼９５，９６の個数を増やせば、この機体９０ｃの横揺れに対する抵抗力は増大する。但し、飛行機等にこの機体９０ｃを使用する場合、この機体９０ｃの外側の筒状安定翼９１には、実験とは違って内外両側に風が流れるので、式１２-（２）中のＨは、式１７-（１）の様になる。 If the number of the radial stabilizer blades 93 and 94 and the number of the cylindrical stabilizer blades 95 and 96 of the airframe 90c are increased, the resistance force against the roll of the airframe 90c increases. However, when this machine body 90c is used for an airplane or the like, the wind flows on the inner and outer sides of the cylindrical stabilizer wing 91 outside the machine body 90c, unlike the experiment. Therefore, H in Expression 12- (2) is Equation 17- (1) is obtained.
Figure JPOXMLDOC01-appb-M000076
Figure JPOXMLDOC01-appb-M000076
　以上の様な発明または発明原理を利用すれば、機体の安定翼に掛かる風圧自体により、その機体またはその機体を備えた飛行機の姿勢制御が自動的に行われるので、姿勢制御のための精密で高感度なセンサーおよび高価で高速度なコンピューターシステムが不必要となる。その上、機体の安定翼の周囲に流れる風流による慣性モーメントの増大効果により、その機体自身またはその機体を備えた飛行機自体に、外部風による影響に対する抵抗力が発生し、その結果、その機体自身またはその機体を備えた飛行機は、外部風に対して非常に強くなる。 If the invention or principle of the invention as described above is used, the attitude control of the aircraft or the airplane equipped with the aircraft is automatically performed by the wind pressure applied to the stable wings of the aircraft. Sensitive sensors and expensive, high-speed computer systems are unnecessary. In addition, due to the effect of increasing the moment of inertia due to the wind flow around the stable wing of the fuselage, the aircraft itself or the airplane equipped with the aircraft itself is resistant to the effects of external wind, and as a result, the aircraft itself Or an airplane with that airframe will be very strong against external winds.
　この発明は、プロペラ機のみではなくジェット機、ロケットおよびガス噴射等の飛行機に対しても応用できるので、広く姿勢制御の必要な分野に利用できる。またこの発明を利用した図１２，図１３または図１４の様なフライングユニット（姿勢制御装置）により、フライングカーの実現も夢ではなくなる。 Since the present invention can be applied not only to propeller aircraft but also to airplanes such as jet aircraft, rockets, and gas injection, it can be widely used in fields requiring attitude control. In addition, with the flying unit (attitude control device) as shown in FIG. 12, 13 or 14 using the present invention, the realization of a flying car is no longer a dream.
　Ｓ１９（プロペラ回転軸の振れによる影響を相殺して機体を安定化させる条件）
　図１５は、例えば図１２の機体８０において、プロペラ回転時のある一瞬を側面から視たときの図である。図１５中、点Ｏは、プロペラ回転軸の固定点であり、点Ｐは、プロペラの作用点（中心点）Ｙを含む水平線と点Ｏを含む垂直線との交点であり、点Ｇ1は、機体８０の重心Ｇが点Ｏよりも下にある場合の重心の位置であり、点Ｇ2は、機体８０の重心Ｇが各点Ｏ，Ｐ間にある場合の重心の位置であり、点Ｇ3は、機体８０の重心Ｇが点Ｐよりも上にある場合の重心の位置である。また点Ｏと機体８０の重心Ｇとの間の距離のｎ値をｎGとする。 S19 (Conditions to stabilize the aircraft by offsetting the influence of the propeller rotating shaft runout)
FIG. 15 is a view of a certain moment when the propeller rotates, for example, in the airframe 80 of FIG. 12 when viewed from the side. In FIG. 15, point O is a fixed point of the propeller rotation axis, point P is an intersection of a horizontal line including the propeller operating point (center point) Y and a vertical line including point O, and point G 1 is , The center of gravity when the center of gravity G of the body 80 is below the point O, and the point G 2 is the position of the center of gravity when the center of gravity G of the body 80 is between the points O and P. G 3 is the position of the center of gravity when the center of gravity G of the body 80 is above the point P. Further, the n value of the distance between the point O and the center of gravity G of the body 80 is n G.
　ここで、機体８０の重心Ｇが各点Ｇ1～Ｇ3にある場合のモーメントバランス式を考える。 Here, consider a moment balance equation when the center of gravity G of the airframe 80 is at each of the points G 1 to G 3 .
　（１）機体８０の重心Ｇが点Ｇ1にある場合
　この場合は、プロペラ８１の推進力ＦPの水平成分（Ｆpsinβ）のモーメントの方が、ＦPの垂直成分（ｍｇ）のモーメントよりも大きくなり、それらのモーメントは互いに逆向きになるので、モーメントバランス式は、式１９-（６）となる。 (1) the moment of this case when the center of gravity G of the machine body 80 is at point G 1, towards the moment of the horizontal component of the thrust F P of the propeller 81 (F p sin .beta) is the vertical component of the F P (mg) Since the moments are opposite to each other, the moment balance equation is expressed by Equation 19- (6).
Figure JPOXMLDOC01-appb-M000077
Figure JPOXMLDOC01-appb-M000077
　尚、式中のＦCは、風圧中心点Ｃに掛かるプロペラ後流の風圧力であり、ＦWは、外部風圧点Ｗに掛かる外部風圧力である。 In the equation, F C is the wind pressure behind the propeller applied to the wind pressure central point C, and F W is the external wind pressure applied to the external wind pressure point W.
　ＦP＝ｍｇ／cosβ、ＦC＝Ｈπｍｇsinβ、ＦW＝Ｗｍｇsinβなので、これらの式を式１９-（６）に代入すると、式１９-（１）を得る。尚、βは、微小な角度なので、cosβ≒１と考えられる。 Since F P = mg / cosβ, F C = Hπmgsinβ, and F W = Wmgsinβ, substituting these equations into Equation 19- (6) yields Equation 19- (1). Since β is a minute angle, it is considered that cos β≈1.
Figure JPOXMLDOC01-appb-M000078
Figure JPOXMLDOC01-appb-M000078
　（２）機体８０の重心Ｇが点Ｇ2にある場合
　この場合は、プロペラ推進力ＦPの水平成分（Ｆpsinβ）と垂直成分（ｍｇ）との各々のモーメントの大きさが上記（１）の場合と逆になるが、モーメントバランス式を考えると、上記（１）と同様に、式１９-（１）を得る。 (2) When the center of gravity G of the airframe 80 is at the point G 2 In this case, the magnitudes of the moments of the horizontal component (F p sin β) and the vertical component (mg) of the propeller propulsive force F P are the above (1 However, considering the moment balance equation, Equation 19- (1) is obtained in the same manner as (1) above.
　（３）機体８０の重心Ｇが点Ｇ3にある場合
　この場合は、プロペラ推進力ＦPの水平成分（Ｆpsinβ）と垂直成分（ｍｇ）との各々のモーメントの向きは同じになるので、上記（１）の場合と同様に、式１９-（１）を得る。 (3) When the center of gravity G of the airframe 80 is at the point G 3 In this case, the directions of the moments of the horizontal component (F p sinβ) and the vertical component (mg) of the propeller thrust force F P are the same. As in the case of the above (1), Expression 19- (1) is obtained.
　上記（１）～（３）より式１９-（１）が常に成立することが分かる。ここで、プロペラ回転軸の振れによる影響を相殺して機体８０を安定化させるには、機体８０に作用するモーメントが釣り合うか、或いはプロペラ回転軸を垂直方向に向けるモーメントが優勢であれば良い（即ち、式１９-（１）の左辺がｎGの値と同じか、或いはそれよりも大きければ良い）ので、プロペラ回転軸の振れによる影響を相殺して機体８０を安定化させるための条件は、式１９-（２）となる。 From the above (1) to (3), it can be seen that Expression 19- (1) always holds. Here, in order to offset the influence of the swing of the propeller rotation shaft and stabilize the airframe 80, the moment acting on the airframe 80 may be balanced, or the moment that directs the propeller rotation shaft in the vertical direction may be dominant ( That is, since the left side of Equation 19- (1) should be equal to or larger than the value of n G ), the condition for stabilizing the body 80 by offsetting the influence of the propeller rotation shaft shake is: Equation 19- (2) is obtained.
Figure JPOXMLDOC01-appb-M000079
Figure JPOXMLDOC01-appb-M000079
　式１９-（１）の右辺のｎGをＳ５（疑似揚力係数ｋの決定）の項で説明したｎXで表すと、式１９-（３）となる。 When n G on the right side of Expression 19- (1) is represented by n X described in the section of S5 (determination of the pseudo lift coefficient k), Expression 19- (3) is obtained.
Figure JPOXMLDOC01-appb-M000080
Figure JPOXMLDOC01-appb-M000080
　式１９-（３）を式１９-（１）に代入し、且つｎGW＝０とおくと、式７-（１９）および式７-（２０）が得られる。よって、式７-（１９）および式７-（２０）は、式１９-（２）の条件の一態様と言える。またｎGCおよびｎGWは、必ず式１９-（４）を満たさなければならない。  By substituting Equation 19- (3) into Equation 19- (1) and setting n GW = 0, Equation 7- (19) and Equation 7- (20) are obtained. Therefore, it can be said that Formula 7- (19) and Formula 7- (20) are one mode of the condition of Formula 19- (2). Also, n GC and n GW must always satisfy the equation 19- (4).
Figure JPOXMLDOC01-appb-M000081
Figure JPOXMLDOC01-appb-M000081
　よって、機体８０が安定してホバリングする条件は、式１９-（５）の様に表す事ができる。 Therefore, the conditions for the airframe 80 to stably hover can be expressed as in Equation 19- (5).
Figure JPOXMLDOC01-appb-M000082
Figure JPOXMLDOC01-appb-M000082
　機体８０が安定してホバリングしているときは、式１９-（５）が必ず成立していると言える。 When the aircraft 80 is stably hovering, it can be said that Equation 19- (5) is always satisfied.
　但し、機体８０の重心Ｇがプロペラ８１よりも上にある場合は、重心Ｇの復元効果が無いばかりでなく、逆に重力Ｇにより機体８０を傾けるモーメントが発生するので、非常に不安定になる。また機体８０の重心Ｇがプロペラ８１よりも下にある場合でも、プロペラ８１から重心Ｇまでの距離ｎPGが短い場合は、重力Ｇによる復元効果が小さいので、プロペラ回転軸の振れによる影響以外の雑力の影響が相対的に大きくなり、機体８０が左右に揺れる可能性が大きくなる。よって、雑力の影響を無視できる程度にｎPGを長く取ることが重要である。ｎPGの長さに制限がある場合は、機体８０の慣性モーメントを大きくするために、式１９-（５）の第２式の左辺をできるだけ大きくすればよい。 However, when the center of gravity G of the fuselage 80 is above the propeller 81, not only is there no effect of restoring the center of gravity G, but conversely, a moment that tilts the fuselage 80 by gravity G is generated, which makes it very unstable. . Even when the center of gravity G of the body 80 is below the propeller 81, if the distance n PG from the propeller 81 to the center of gravity G is short, since the restoring effect of the gravitational force G is small, other than the effect of vibration of the propeller shaft The influence of the miscellaneous force becomes relatively large, and the possibility that the airframe 80 swings left and right is increased. Therefore, it is important to take n PG long enough to ignore the influence of miscellaneous power. When the length of n PG is limited, in order to increase the moment of inertia of the airframe 80, the left side of the second expression of Expression 19- (5) may be increased as much as possible.
　Ｓ２０（筒状安定翼内のプロペラ位置によらない風圧中心点の位置固定）
　例えば図１２の機体８０において、筒状安定翼８３内の放射状安定翼８２の長さを例えば半分位に短くして、放射状安定翼８２と共にプロペラ８１の位置を筒状安定翼８３の鉛直方向に変化させ、そのそれぞれの位置において、機体８０が安定してホバリングする重心位置を求める実験を行った。 S20 (fixed position of wind pressure center point independent of propeller position in cylindrical stabilizer blade)
For example, in the fuselage 80 of FIG. 12, the length of the radial stabilizer wing 82 in the cylindrical stabilizer wing 83 is shortened to, for example, about half, and the position of the propeller 81 together with the radial stabilizer wing 82 is set in the vertical direction of the cylindrical stabilizer wing 83. An experiment was performed in which the center of gravity position at which the airframe 80 stably hovered was obtained at each position.
　その結果、プロペラ位置が筒状安定翼８３の上端から筒状安定翼８３の長さの１／８以上に下に位置したときの実験では、機体８０の各安定翼８２，８３に掛かる総合風圧力が全て、筒状安定翼８３の上端から筒状安定翼８３の長さの１／８だけ下がった点付近に集中して掛かっている事が分かった。プロペラ位置を更に下げた実験においても、機体８０の各安定翼８２，８３に掛かる総合風圧力が集中して掛かる点（総合風圧中心点）の位置は、変化せず、筒状安定翼８３の上端から筒状安定翼８３の長さの１／８だけ下がった点付近のままであった。 As a result, in an experiment in which the propeller position is positioned below the upper end of the cylindrical stabilizer wing 83 to 1/8 or more of the length of the cylindrical stabilizer wing 83, the total wind applied to the respective stable wings 82 and 83 of the fuselage 80 is determined. It was found that all of the pressure was concentrated in the vicinity of a point that was lowered from the upper end of the cylindrical stabilizing blade 83 by ８ of the length of the cylindrical stabilizing blade 83. Even in the experiment in which the propeller position is further lowered, the position of the point where the total wind pressure applied to each of the stable blades 82 and 83 of the fuselage 80 is concentrated (the total wind pressure central point) does not change, and the cylindrical stable blade 83 It remained in the vicinity of a point lowered by 1/8 of the length of the cylindrical stabilizing blade 83 from the upper end.
　言い換えれば、プロペラ位置が筒状安定翼８３の上端から筒状安定翼８３の長さの１／８以上に下がると、機体８０の総合風圧中心点が筒状安定翼８３の上端から筒状安定翼８３の長さの１／８だけ下がった点に固定されるということである。尚、プロペラ位置が筒状安定翼８３の上端から筒状安定翼８３の長さの１／８だけ下がった点よりも上に位置する場合は、Ｓ１３で説明した通りの位置に、機体８０の総合風圧中心点が現れていた。 In other words, when the propeller position is lowered from the upper end of the cylindrical stabilizer wing 83 to 1/8 or more of the length of the cylindrical stabilizer wing 83, the total wind pressure center point of the airframe 80 is stabilized from the upper end of the cylindrical stabilizer wing 83. That is, it is fixed at a point lowered by 1/8 of the length of the wing 83. When the propeller position is above the point where the length of the cylindrical stabilizer wing 83 is lowered from the upper end of the cylindrical stabilizer wing 83 by one-eighth, the position of the airframe 80 is changed to the position described in S13. The total wind pressure center point appeared.
　尚、上記の実験において機体８０の総合風圧中心点の位置は、以下の様に求めた。即ち、図１６の様に、機体８０の筒状安定翼８３の下方に連結部８３ｃを介して筒状の補助安定翼８３ｂを同心軸状に取り付けた機体８０ａを作成し、安定してホバリングする機体８０ａの重心を求めることで、機体８０の総合風圧中心点を求められる。機体８０に掛かる総合風圧力および補助安定翼８３ｂに掛かる風圧力の各々の大きさの比は計算できる。そして外部風に対する風圧中心点の位置も計算できるので、式１９-（５）の第１式（ＨπｎGC＝ＷｎGW）よりｎGCが求まり、機体８０ａの各安定翼８２，８３，８３ｂに掛かる総合風圧力が集中して掛かる点（総合風圧中心点）の位置が決定される。その後、機体８０全体に掛かる風圧力と補助安定翼８３ｂに掛かる風圧力との比に基づき、機体８０の総合風圧中心点の位置を求める事ができる。 In the above experiment, the position of the total wind pressure center point of the airframe 80 was obtained as follows. That is, as shown in FIG. 16, a machine body 80a in which a cylindrical auxiliary stabilizer blade 83b is attached concentrically through a connecting portion 83c below the cylindrical stabilizer blade 83 of the machine body 80 is stably hovered. By obtaining the center of gravity of the body 80a, the total wind pressure center point of the body 80 can be obtained. The ratio of the magnitudes of the total wind pressure applied to the airframe 80 and the wind pressure applied to the auxiliary stabilizing blade 83b can be calculated. Since the position of the wind pressure center point with respect to the external wind can also be calculated, n GC is obtained from the first equation (Hπn GC = Wn GW ) of Equation 19- (5) and applied to each of the stable blades 82, 83, 83b of the airframe 80a. The position of the point at which the total wind pressure is concentrated (the total wind pressure central point) is determined. Thereafter, based on the ratio of the wind pressure applied to the entire body 80 and the wind pressure applied to the auxiliary stabilizing blade 83b, the position of the central wind pressure center point of the body 80 can be obtained.
　上記の実験の結果は、予想もできない結果となった。このことは、機体８０ａにおいて、各安定翼８３，８３ｂのサイズおよび各安定翼８３，８３ｂ間の距離を全く変えずに、放射状安定翼８２と共にプロペラ８１の位置を上下に変化させても、安定してホバリングする重心位置が変化しなかったことからも分かる。 The result of the above experiment was an unexpected result. This means that in the fuselage 80a, even if the position of the propeller 81 is changed up and down together with the radial stabilizer 82 without changing the size of the stabilizers 83 and 83b and the distance between the stabilizers 83 and 83b at all, It can also be seen from the fact that the position of the center of gravity for hovering did not change.
　この事実は、機体を設計の自由度を向上させる上で大変都合が良い。プロペラ位置を筒状安定翼８３の上端から筒状安定翼８３の長さの１／８以上下がった位置に配置すれば、放射状安定翼８２の位置は、任意に選択できる事になる。但し、このとき、機体８０全体を安定させるためには、図１６の様に、機体８０全体の下に、筒状または放射状の補助安定翼８３ｂを取り付ける必要がある。 This fact is very convenient for improving the degree of freedom in designing the aircraft. If the propeller position is disposed at a position lower than 1/8 or more of the length of the cylindrical stabilizer blade 83 from the upper end of the cylindrical stabilizer blade 83, the position of the radial stabilizer blade 82 can be arbitrarily selected. However, at this time, in order to stabilize the entire body 80, it is necessary to attach a cylindrical or radial auxiliary stabilizing wing 83b under the entire body 80 as shown in FIG.
　この場合、補助安定翼８３ｂは機体８０の下ではなく上に取り付けても良いが、上に取り付けた場合は、補助安定翼８３ｂに掛かる風圧力は、下に取り付けた場合に補助安定翼８３ｂに掛かる風圧力に比べて非常に弱いため、補助安定翼８３ｂの投影面積を非常に大きくするか、或いは、機体８０から非常に離れた位置に取り付ける必要があるので、現実的ではない。 In this case, the auxiliary stabilizer wing 83b may be attached to the auxiliary stabilizer wing 83b instead of below the fuselage 80. However, if the auxiliary stabilizer wing 83b is attached to the top, the wind pressure applied to the auxiliary stabilizer wing 83b Since it is very weak compared to the wind pressure applied, it is not realistic because the projected area of the auxiliary stabilizing blade 83b needs to be very large or mounted at a position very far from the airframe 80.
　図１６の機体８０ａの様に、筒状安定翼８３の下方に補助安定翼８３ｂを取り付けると、その補助安定翼８３ｂに掛かる風圧力のために、機体８０ａ全体の総合風圧中心点Ｃ0が結果的に下に下がることになる。そのため、機体８０ａを安定してホバリングさせるための重心位置も下に下がることになり、重心による復元効果が増大する。その上、機体８０ａ全体の慣性モーメントも同じく増大するので、機体８０ａは、補助安定翼８３ｂが無い場合に比べてより一層に安定することになる。 When the auxiliary stabilizer wing 83b is attached below the cylindrical stabilizer wing 83 as in the fuselage 80a of FIG. 16, the total wind pressure center point C 0 of the entire fuselage 80a is the result of the wind pressure applied to the auxiliary stabilizer wing 83b. Will go down. Therefore, the position of the center of gravity for stably hovering the airframe 80a is also lowered, and the restoring effect by the center of gravity is increased. In addition, since the inertia moment of the entire body 80a is also increased, the body 80a is further stabilized as compared with the case where the auxiliary stabilizing wing 83b is not provided.
　尚、上記の実施の形態１～３の機体では、風流を発生させる手段（風流発生装置）を、プロペラとプロペラ用駆動部とにより構成したが、その様に限定するものではなく、例えばガス噴射機、ジェット噴流機またはロケット噴射機により構成しても良い。風流発生装置をプロペラとプロペラ用駆動部とにより構成すれば、比較的簡単な原理で風流を発生させる事ができる。また風流発生装置をガス噴射機、ジェット噴流機またはロケット噴射機により構成すれば、より強力な風流を発生させる事ができる。 In the airframes of the first to third embodiments described above, the means for generating the wind flow (wind flow generating device) is constituted by the propeller and the propeller drive unit. However, the present invention is not limited to this, for example, gas injection You may comprise by a machine, a jet jet machine, or a rocket jet machine. If the wind flow generating device is composed of a propeller and a propeller drive unit, it is possible to generate a wind flow on a relatively simple principle. Further, if the wind flow generating device is constituted by a gas jet, jet jet or rocket jet, a stronger wind flow can be generated.

Claims (17)

  1. 　放射状に組まれた複数の垂直主翼を備え、
    　プロペラの直径をｒ0としたときに、前記プロペラが周囲の空気から受けるモーメントの２πｒ0倍のモーメントが、プロペラ後流により前記複数の垂直主翼に掛かるように、前記垂直主翼の形状が形成されることを特徴とするプロペラ機。     With multiple vertical wings assembled radially,
    The shape of the vertical main wing is formed such that when the propeller diameter is r 0 , a moment 2πr 0 times the moment that the propeller receives from the surrounding air is applied to the plurality of vertical main wings by the propeller wake. Propeller aircraft characterized by that.
  2. 　前記機体は、放射状に組まれた複数の垂直主翼を備え、
    　前記各垂直主翼は、プロペラ後流の拡がりと同じ広さに形成され、
    　前記各垂直主翼の高さΔｌは、前記プロペラの直径をｒ0とし、前記垂直主翼の直径単位の枚数をｍとすると、Δｌ＝２πｒ0 ／ｍを満たすように設定されることを特徴とする請求項１に記載のプロペラ機。     The airframe includes a plurality of vertical wings assembled radially,
    Each of the vertical main wings is formed in the same area as the spread of the propeller wake,
    The height Δl of each vertical main wing is set to satisfy Δl = 2πr 0 / m, where r 0 is the diameter of the propeller and m is the number of units of the diameter of the vertical main wing. The propeller machine according to claim 1.
  3. 　放射状に組まれた複数枚の垂直仕切板を有する機体と、
    　前記複数枚の垂直仕切板の上側に配置されたプロペラと、
    を備え、
    　前記プロペラの直径をｒ0としたときに、前記プロペラが周囲の空気から受けるモーメントの２πｒ0倍のモーメントが、プロペラ後流により前記複数の垂直仕切板に掛かるように、前記垂直仕切板の形状が形成されることを特徴とするプロペラ装置。     A fuselage having a plurality of vertical partition plates assembled radially;
    A propeller disposed above the plurality of vertical partition plates;
    With
    The shape of the vertical partition plate is such that when the diameter of the propeller is r 0 , a moment 2πr 0 times the moment that the propeller receives from the surrounding air is applied to the plurality of vertical partition plates by the wake of the propeller. A propeller device characterized in that is formed.
  4. 　前記各垂直仕切板は、プロペラ後流の拡がりと同じ広さに形成され、
    　前記各垂直主翼の高さΔｌは、前記プロペラの直径をｒ0とし、前記垂直仕切板の直径単位の枚数をｍとすると、Δｌ＝２πｒ0 ／ｍを満たすように設定されることを特徴とする請求項３に記載のプロペラ装置。     Each of the vertical partition plates is formed in the same area as the spread of the propeller wake,
    The height Δl of each vertical main wing is set to satisfy Δl = 2πr 0 / m, where r 0 is the diameter of the propeller and m is the number of diameter units of the vertical partition plate. The propeller device according to claim 3.
  5. 　前記機体は、前記複数枚の垂直仕切板の周囲に被された筒状体を更に有することを特徴とする請求項３または４に記載のプロペラ装置。 The propeller device according to claim 3 or 4, wherein the airframe further includes a cylindrical body covered around the plurality of vertical partition plates.
  6. 　前記機体は、ｎ（ｎ：２以上の整数）層に分割構成され、それら各層はそれぞれ、放射状に組まれた複数枚の垂直仕切板を備え、
    　ｉ番目の前記層の高さおよび垂直仕切板の直径単位の枚数をそれぞれΔｌiおよびｍiとすると、前記各層の高さΔｌ1，Δｌ2，，…，Δｌnおよび垂直仕切板の直径単位の枚数ｍ1，ｍ2，，…，ｍnは、Δｌ1ｍ1＋Δｌ2ｍ2＋…＋Δｌnｍn≒２πｒ0を満たすように設定されることを特徴とする請求項３に記載のプロペラ装置。     The aircraft is configured to be divided into n (n: an integer of 2 or more) layers, each of which includes a plurality of vertical partition plates that are radially assembled,
    When i-th of the height of the layer and the number of diameter unit of a vertical partition plate respectively and .DELTA.l i and m i, height .DELTA.l 1 of each layer, Δl 2 ,, ..., Δl n and diameter unit of a vertical partition plate 4. The number m 1 , m 2 ,..., Mn of the first and second sheets is set to satisfy Δl 1 m 1 + Δl 2 m 2 +... + Δl n m n ≈2πr 0 . Propeller device.
  7. 　前記各層はそれぞれ、前記複数枚の垂直仕切板の周囲に被された筒状体を更に備えることを特徴とする請求項６に記載のプロペラ装置。 The propeller device according to claim 6, wherein each of the layers further includes a cylindrical body covered around the plurality of vertical partition plates.
  8. 　放射状安定翼または筒状安定翼を備えることを特徴とする姿勢制御装置。 An attitude control device comprising radial stabilizers or cylindrical stabilizers.
  9. 　筒状安定翼と、
    　前記筒状安定翼の中心軸線上に沿って同軸線状に配設された１つ以上の放射状安定翼と、
    を備えることを特徴とする姿勢制御装置。     A cylindrical stabilizer,
    One or more radial stabilizers disposed coaxially along the central axis of the cylindrical stabilizer;
    An attitude control device comprising:
  10. 　前記筒状安定翼と、その内部に同軸線状に１つ以上の筒内筒状安定翼とを更に備えることを特徴とする請求項８または９に記載の姿勢制御装置。 The attitude control device according to claim 8 or 9, further comprising: the cylindrical stabilizer wing and one or more in-cylinder cylindrical stabilizer wings coaxially therein.
  11. 　前記筒状安定翼または／および前記放射状安定翼の中心軸線上に配設され、前記中心軸線方向に風流を発生させる風流発生装置を更に備えることを特徴とする請求項８～１０の何れかに記載の姿勢制御装置。 11. The wind flow generator according to claim 8, further comprising a wind flow generator disposed on a central axis of the cylindrical stabilizer or / and the radial stabilizer, and generating a wind flow in the central axis direction. The attitude control device described.
  12. 　前記放射状安定翼または／および前記筒状安定翼は、当該姿勢制御装置の重心と各安定翼の風圧中心点の総合風圧中心点との距離ｎGCおよび重心と外部風圧中心点Ｗとの距離ｎGWとの関係が式１９-（５）で表される様に配置されることを特徴とする請求項８～１１の何れかに記載の姿勢制御装置。
    Figure JPOXMLDOC01-appb-M000001
         The radial stabilizer wing or / and the cylindrical stabilizer wing are the distance n GC between the center of gravity of the attitude control device and the total wind pressure center point of the wind pressure center point of each stabilizer blade, and the distance n between the center of gravity and the external wind pressure center point W. The attitude control device according to any one of claims 8 to 11, wherein the attitude control device is arranged so that a relationship with GW is expressed by an expression 19- (5).
    Figure JPOXMLDOC01-appb-M000001
  13. 　前記筒状安定翼の上端から下に前記筒状安定翼の長さの１／８以上の距離の位置に、前記筒状安定翼の中心軸線方向に風流を発生させる風流発生装置を配置したことを特徴とする請求項１１または１２に記載の姿勢制御装置。 An airflow generator for generating an airflow in the direction of the central axis of the cylindrical stabilizer blade is disposed at a position of a distance of 1/8 or more of the length of the cylindrical stabilizer blade from the upper end of the cylindrical stabilizer blade. The attitude control device according to claim 11 or 12,
  14. 　前記風流発生装置は、風流を発生させるプロペラと、プロペラ用駆動部とを備えることを特徴とする請求項１１または１３に記載の姿勢制御装置。 The attitude control device according to claim 11 or 13, wherein the wind flow generation device includes a propeller that generates a wind flow and a propeller drive unit.
  15. 　前記筒状安定翼の下に補助安定翼を配置させたことを特徴とする請求項１１～１３の何れかに記載の姿勢制御装置。 The attitude control device according to any one of claims 11 to 13, wherein an auxiliary stabilizer blade is disposed below the cylindrical stabilizer blade.
  16. 　前記放射状安定翼が中心軸線対称に形成された場合において、
    　前記プロペラの実効角度をβTとし、ｉ番目の前記放射状安定翼の安定翼の直径単位の枚数をｍiとし、ｉ番目の前記放射状安定翼の中心軸線方向の長さを前記プロペラの直径で割った値をｎiとすると、式１４-（１）が成立することを特徴とする請求項１１～１５の何れかに記載の姿勢制御装置。
    Figure JPOXMLDOC01-appb-M000002
         In the case where the radial stabilizer is formed symmetrically about the central axis,
    And the effective angle beta T of the propeller, the number of diameters units of the i-th of said radial stabilizing wings stable wing and m i, the i-th center axis direction of the length of the radial stabilizing wing diameter of the propeller The attitude control device according to any one of claims 11 to 15, wherein equation 14- (1) is established, where n i is a divided value.
    Figure JPOXMLDOC01-appb-M000002
  17. 　請求項１１～１６の何れかの姿勢制御装置のうちの同じものを２つ組み合わせた姿勢制御装置であって、
    　それぞれその吸気側開口端を上側に向けると共にその排気側開口端を下側に向け、且つ互いの吸気側開口端を互いの対向方向に傾斜（傾斜角度０°も含む）させる様にして、互いに間隔空けて配置された前記２つの姿勢制御装置と、
    　前記２つの姿勢制御装置を相互連結する連結部材と、
    を備えることを特徴とする姿勢制御装置。     A posture control device combining two of the same posture control devices of any one of claims 11 to 16,
    Respectively, the intake side open end is directed upward, the exhaust side open end is directed downward, and the intake side open ends are inclined in directions opposite to each other (including an inclination angle of 0 °). The two attitude control devices arranged at an interval;
    A connecting member that interconnects the two attitude control devices;
    An attitude control device comprising:
PCT/JP2008/051415 2008-01-30 2008-01-30 Propeller aircraft, propeller apparatus, and posture controller WO2009096010A1 (en)

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