WO2009096010A1 - Propeller aircraft, propeller apparatus, and posture controller - Google Patents
Propeller aircraft, propeller apparatus, and posture controller Download PDFInfo
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- 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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- propeller
- stabilizer
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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
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.
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>
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。 <
In this embodiment, the conditions for stabilizing the hovering of the
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.
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).
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).
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).
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).
そしてrl=2ωRなので、式(6)となる。
Since r l = 2ωR, equation (6) is obtained.
よって式(7)を得る。
Therefore, formula (7) is obtained.
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).
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).
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).
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).
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).
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 themain wing 41a matches the instantaneous moment of the propeller 43.
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
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.
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).
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.
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.
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).
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.
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).
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).
そしてMac=Mcbから、式(64)が成立する。
And from Mac = Mcb, Formula (64) is materialized.
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, thepropeller 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.
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。 <
As shown in FIG. 5, the
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.
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.
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。 <
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 atriangular 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.
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
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 thepropeller 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
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
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.
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).
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.
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.
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.
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.
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).
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).
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 stabilizingwing 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).
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
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). .
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).
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.
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.
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.
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 thepropeller 60 to the wind pressure center point C due to the wake of the propeller is expressed by Expression 1- (1).
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
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.
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 thefuselage 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.
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
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 theairframe 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.
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
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.
またFCの定義式は式3-(1)である。
The defining formula of F C is Formula 3- (1).
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.
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 theairframe 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.
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
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.
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).
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).
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.
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.
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 stabilizingblade 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.
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
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.
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.
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) ′.
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 anaircraft 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.
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
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.
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).
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.
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.
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).
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)
Theairframe 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
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).
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?
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.
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).
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.
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).
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).
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.
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.
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).
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).
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).
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)
Theairframe 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).
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
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. .
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).
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 acylindrical 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.
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
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).
(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
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.
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 twoairframes 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.
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
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.
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).
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.
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).
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).
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).
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.
(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
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).
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.
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 theairframe 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.
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
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 themachine 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).
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
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.
(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 theairframe 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.
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
(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 theairframe 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.
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
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.
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.
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).
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).
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 thefuselage 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.
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
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)
- 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. - 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. - 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. - 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. - 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.
- 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. - 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.
- 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.
- 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: - 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.
- 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.
- 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。
- 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,
- 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.
- 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.
- 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。
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.
- 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:
Priority Applications (4)
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PCT/JP2008/051415 WO2009096010A1 (en) | 2008-01-30 | 2008-01-30 | Propeller aircraft, propeller apparatus, and posture controller |
PCT/JP2008/059529 WO2009096048A1 (en) | 2008-01-30 | 2008-05-23 | Propeller aircraft, propeller apparatus, posture controller, propelling force amplifier, and flight equipment |
JP2009551392A JP5184555B2 (en) | 2008-01-30 | 2008-09-03 | Propeller device, attitude control device, propulsion force amplifying device, and flying device |
PCT/JP2008/065873 WO2009096058A1 (en) | 2008-01-30 | 2008-09-03 | Propeller plane, propeller device, posture controller, propulsive force amplification device and flight device |
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PCT/JP2008/051415 WO2009096010A1 (en) | 2008-01-30 | 2008-01-30 | Propeller aircraft, propeller apparatus, and posture controller |
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PCT/JP2008/051415 WO2009096010A1 (en) | 2008-01-30 | 2008-01-30 | Propeller aircraft, propeller apparatus, and posture controller |
PCT/JP2008/059529 WO2009096048A1 (en) | 2008-01-30 | 2008-05-23 | Propeller aircraft, propeller apparatus, posture controller, propelling force amplifier, and flight equipment |
PCT/JP2008/065873 WO2009096058A1 (en) | 2008-01-30 | 2008-09-03 | Propeller plane, propeller device, posture controller, propulsive force amplification device and flight device |
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PCT/JP2008/059529 WO2009096048A1 (en) | 2008-01-30 | 2008-05-23 | Propeller aircraft, propeller apparatus, posture controller, propelling force amplifier, and flight equipment |
PCT/JP2008/065873 WO2009096058A1 (en) | 2008-01-30 | 2008-09-03 | Propeller plane, propeller device, posture controller, propulsive force amplification device and flight device |
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Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2012014272A1 (en) * | 2010-07-26 | 2012-02-02 | 川口 泰子 | Flight vehicle |
US9016616B2 (en) | 2010-07-26 | 2015-04-28 | Hiroshi Kawaguchi | Flying object |
Families Citing this family (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20210370733A1 (en) * | 2019-12-23 | 2021-12-02 | California Institute Of Technology | Synchronized Multi-Modal Robot |
KR102485309B1 (en) * | 2022-06-16 | 2023-01-06 | 이춘형 | Flying car capable of vertical take-off and landing |
Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPH06293296A (en) * | 1992-12-28 | 1994-10-21 | Hughes Missile Syst Co | Pilotless aircraft for effecting vertical take off and landing and level cruise flight |
US5746390A (en) * | 1996-03-20 | 1998-05-05 | Fran Rich Chi Associates, Inc. | Air-land vehicle with ducted fan vanes providing improved performance |
JPH11217099A (en) * | 1998-01-30 | 1999-08-10 | Baitekkusu:Kk | Aero-carrier |
Family Cites Families (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US5419514A (en) * | 1993-11-15 | 1995-05-30 | Duncan; Terry A. | VTOL aircraft control method |
AU2251500A (en) * | 1998-08-27 | 2000-04-03 | Nicolae Bostan | Gyrostabilized self propelled aircraft |
US7658346B2 (en) * | 2005-02-25 | 2010-02-09 | Honeywell International Inc. | Double ducted hovering air-vehicle |
-
2008
- 2008-01-30 WO PCT/JP2008/051415 patent/WO2009096010A1/en active Application Filing
- 2008-05-23 WO PCT/JP2008/059529 patent/WO2009096048A1/en active Application Filing
- 2008-09-03 WO PCT/JP2008/065873 patent/WO2009096058A1/en active Application Filing
Patent Citations (3)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
JPH06293296A (en) * | 1992-12-28 | 1994-10-21 | Hughes Missile Syst Co | Pilotless aircraft for effecting vertical take off and landing and level cruise flight |
US5746390A (en) * | 1996-03-20 | 1998-05-05 | Fran Rich Chi Associates, Inc. | Air-land vehicle with ducted fan vanes providing improved performance |
JPH11217099A (en) * | 1998-01-30 | 1999-08-10 | Baitekkusu:Kk | Aero-carrier |
Cited By (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
WO2012014272A1 (en) * | 2010-07-26 | 2012-02-02 | 川口 泰子 | Flight vehicle |
US9016616B2 (en) | 2010-07-26 | 2015-04-28 | Hiroshi Kawaguchi | Flying object |
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