JP5184555B2 - Propeller device, attitude control device, propulsion force amplifying device, and flying device - Google Patents

Description

  The present invention relates to a propeller aircraft capable of stable vertical take-off and landing and hovering and application of the technology.

  Currently, there are helicopters and similar airplanes (for example, a tilt rotor type US military aircraft V-22) as the aircraft that can take off and landing in the air and in the air. Other than that, the 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.

  Now, in this situation, if a method for solving the above problems is discovered, dramatic development can be expected in the field of aviation.

  Note that Patent Documents 1 to 8 have already been published as prior art documents related to the background art of the present invention.

Japanese Utility Model Publication No. 4-5199 JP-A-6-293296 JP 2006-327219 A JP 2007-118891 A JP 2005-533700 A Special table 2007-521174 gazette JP-A-5-39092 Japanese Unexamined Patent Publication No. 7-232699

  The above-described practical machines developed so far have a 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.

  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.

  The problem of controlling the posture of an object flying in the sky was a very difficult task because there was no object to support itself around it, and it stood in front of mankind since the beginning of mankind.

  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 big burden was applied.

  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.

A propeller device according to a first aspect of the present invention includes an airframe having a plurality of vertically divided 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.
Further, in the propeller device, the airframe is divided into n (n: an integer of 2 or more) layers, and each of the layers includes a plurality of vertically divided vertical partition plates, and the i-th layer When the height 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 the 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.
A propeller device according to a second aspect of the present invention includes a fuselage 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.

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 wake of the propeller, 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.
Furthermore, 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.

In the propeller device according to the second aspect of the present invention, each of the layers further includes a cylindrical body covered around the plurality of vertical partition plates.

  According to this aspect, the propeller wind can be effectively ejected to the rear of the propeller.

The attitude control device according to the third aspect of the present invention includes a radial stabilizer or a cylindrical stabilizer.
Further, the attitude control device is configured such that the radial stabilizer blade and / or the cylindrical stabilizer blade is a distance n _GC between the gravity center of the attitude control device and the total wind pressure center point of the wind pressure center point of each stabilizer blade and the gravity center and the external The relationship with the distance n _{GW from} the wind pressure center point W is arranged as represented by Expression 19- (5).

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.
Furthermore, according to this aspect, the attitude control device can be stably hovered.

An attitude control device according to a fourth 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.
Further, the attitude control device is configured such that the radial stabilizer blade and / or the cylindrical stabilizer blade is a distance n _GC between the gravity center of the attitude control device and the total wind pressure center point of the wind pressure center point of each stabilizer blade and the gravity center and the external The relationship with the distance n _{GW from} the wind pressure center point W is arranged as represented by Expression 19- (5).

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.
Furthermore, according to this aspect, the attitude control device can be stably hovered.

The attitude control device according to the fifth aspect of the present invention is the central axis of the cylindrical stabilizer blade at a position 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.

  According to this aspect, the position of the radial stabilizer can be arbitrarily selected without changing the position of the center of gravity at which the attitude control device stably hovers, and the degree of freedom in designing the attitude control device can be improved.

In a posture control device according to a sixth aspect of the present invention, the wind flow generator includes a propeller that generates a wind flow and a propeller drive unit.

  According to this embodiment, it is possible to generate a wind flow by a relatively simple principle using a propeller.

In a posture control apparatus according to a seventh aspect of the present invention, auxiliary stabilizer blades are arranged under the cylindrical stabilizer blades.

  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.

In the attitude control device according to the eighth aspect of the present invention, when the radial stabilizer is 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 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.

  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.

The posture control device according to the ninth aspect of the present invention is a posture control device that combines two of the same three of the third to eighth posture control devices, each of which has its intake side open end on the upper side. The two attitude controls arranged at a distance from each other so that the exhaust-side opening end faces downward and the intake-side opening ends face each other at an inclination angle of 0 °. And a connecting member that interconnects the two attitude control devices.

  According to this aspect, the attitude control device can be further stabilized against roll.

An attitude control apparatus according to a tenth aspect of the present invention is an attitude control apparatus in which two or more attitude control apparatuses in which the inclination angle of the attitude control apparatus according to the ninth aspect is greater than 0 ° are combined. Two or more attitude control devices are interconnected so as to cross radially at their central axes in plan view.

  According to this aspect, since two or more attitude control devices are connected radially, it is possible to stabilize against radial rolls and counter-rotations caused by propeller rotation. Further, since the inclination angle is set larger than 0 °, the roll in the radial direction can be further stabilized.

An attitude control apparatus according to an eleventh aspect of the present invention is the attitude control apparatus according to the ninth or tenth aspect, wherein a distance n _GC between the total gravity center G and the total wake wind pressure center point C and the total gravity center thereof. The distance n _GW between G and the total external wind pressure center point W satisfies the equation 19- (5).

  According to this aspect, the attitude control device can be stabilized against roll.

The posture control device according to a twelfth aspect of the present invention is the posture control device according to the tenth aspect, wherein the number of the posture control devices to be combined is an even number of four or more, and its total center of gravity and its total wake wind pressure The center point and its total external wind pressure center point are made to coincide with each other.

  According to this aspect, since the number of attitude control devices to be combined is an even number of 4 or more, the position of the total wake wind pressure center point can be made constant with respect to any radial roll, and after the total Since the total center of gravity and the total external wind pressure center point are matched with the flow wind pressure center point, the total center of gravity, the total wake wind pressure center point, and the total external wind pressure center point can always be matched to any radial roll. As a result, it is possible to strongly stabilize the roll in any radial direction.

The attitude control device according to a thirteenth aspect of the present invention further comprises a stabilizing blade outside the cylindrical stabilizing blade.

  According to this aspect, the attitude control device can be stabilized against rolling and counter-rotation by propeller rotation, and the position of the external wind pressure center point can be adjusted.

An attitude control apparatus according to a fourteenth aspect of the present invention is an attitude control apparatus that combines two or more of the same attitude control apparatuses as described in any one of the third to eighth or thirteenth aspects, The two or more attitude control devices arranged with their axes parallel to each other, with their intake sides facing up and their exhaust sides facing down and spaced apart from each other; and the two or more attitude controls And a connecting member that interconnects the devices.

  According to this aspect, since two or more attitude control devices are connected to each other with the central axes parallel to each other and spaced apart from each other, the attitude control device can be stabilized against rolling and counter-rotation due to propeller rotation. .

The attitude control device according to the fifteenth aspect of the present invention is the attitude control device according to the fourteenth aspect, wherein the total center of gravity G ₀ and the wake wind pressure center point C _i (each of the two or more attitude control devices). i: distance n _{GCi in} the central axis direction with respect to two or more attitude control devices) and a distance n _{GW in the} central axis direction between the total gravity center G ₀ and the total external wind pressure central point W ₀ Satisfies the formula 21- (44) -3.

  According to this aspect, the posture control device can be stabilized against roll and counter-rotation due to propeller rotation.

The posture control device according to the sixteenth aspect of the present invention is the posture control device according to the fourteenth aspect, wherein the total center of gravity G ₀ and the total wake wind pressure center point C ₀ are made to coincide with each other.

  According to this aspect, the attitude control device can be further stabilized against roll.

A propulsive force amplifying device according to a seventeenth aspect of the present invention is a propulsive force amplifying device using, as a propulsion unit, a combination of two or more of the attitude control devices according to fifteenth or sixteenth aspects, The two or more attitude control devices arranged substantially symmetrically so as to direct the propulsive force in the same rotational direction along the coaxial circumference, and a connecting member that interconnects the two or more attitude control devices; Is provided.

  According to this aspect, the propulsive force amplifying device can be levitated with a propulsive force that is a fraction of the gravity of the propulsive force amplifying device, and the propulsive force can be amplified using this.

A propulsive force amplifying device according to an eighteenth aspect of the present invention is the propulsive force amplifying device according to the seventeenth aspect, wherein each of the two or more attitude control devices includes an overall wake wind pressure center point and the coaxial circumference. The straight line connecting the centers of the two and the central axis thereof are interconnected so as to be orthogonal to each other.

  According to this aspect, the propulsive force amplifying device can rotate in a well-balanced manner, and the propulsive force of each attitude control device can be efficiently used for the rotation of the propulsive force amplifying device.

A posture control device according to a nineteenth aspect of the present invention is a posture control device in which a plurality of posture control devices according to any one of the third to twelfth or fourteenth aspects are combined, and its comprehensive propulsive force action point, The total center of gravity, the total wake wind pressure center point, and the total external wind pressure center point are arranged on the center axis.

  According to this aspect, when a plurality of posture control devices are provided, the posture control device can be stabilized against rolling and counter-rotation due to propeller rotation.

A flying device according to a twentieth aspect of the present invention is a flying device using, as a propulsion unit, a combination of two of the same three of the third to eighth attitude control devices, and each of the intake devices. The side opening ends are directed upward, the exhaust side opening ends are directed downward, and the intake side opening ends are opposed to each other at an inclination angle of 0 °. The two attitude control devices and a connecting member that interconnects the two attitude control devices.

  According to this aspect, the flying device can be further stabilized against roll.

A flying device according to a twenty-first aspect of the present invention is a flying device using the attitude control device according to any one of the third to twelfth as a propulsion unit, and further includes a stabilizing wing outside the cylindrical stabilizing wing. It is to be prepared.

  According to this aspect, the flying device can be stabilized against rolling and counter-rotation due to propeller rotation, and the position of the external wind pressure center point can be adjusted.

A flying device according to a twenty-second aspect of the present invention is a flying device using as a propulsion unit a combination of two or more of the same attitude control devices according to any of the third to eighth or thirteenth aspects. The two or more attitude control devices disposed so as to be spaced apart from each other, with their central axes parallel to each other, their intake side facing upward and their exhaust sides facing downward And a connecting member that interconnects two or more attitude control devices.

  According to this aspect, since the two or more attitude control devices are coupled with the central axis being parallel to each other with a space therebetween, the flying device can be stabilized against roll and counter-rotation due to propeller rotation.

Flying device of the 23rd aspect of the present invention is a flying device according to 22, the overall center of gravity G ₀ and the two or more attitude control system each wake wind pressure center point C _i of (i: The distance n _{GCi in} the central axis direction to the identification number of the two or more attitude control devices) and the distance n _{GW in the} central axis direction between the total gravity center G ₀ and the total external wind pressure central point W ₀ The relational expression satisfies Expression 21- (44) -3.

  According to this aspect, the flying device can be stabilized against rolling and counter-rotation due to propeller rotation.

A propulsive force amplifying device according to a twenty-fourth aspect of the present invention is a flying device using the attitude control device according to the fourteenth aspect as a propulsion unit, and has a total center of gravity G ₀ and a total wake wind pressure center point C ₀ . Are matched with each other.

  According to this aspect, the flying device can be further stabilized against rolling.

A flying device according to a twenty-fifth aspect of the present invention is a flying device using a combination of a plurality of attitude control devices according to any of the third to twelfth or fourteenth as a propulsion unit, The propulsive force action point, the total center of gravity, the total wake wind pressure center point, and the total external wind pressure center point are arranged on the center axis.

  According to this aspect, when a plurality of attitude control devices are provided, the flying device can be stabilized against rolling and counter-rotation due to propeller rotation.

A posture control device according to a twenty-sixth aspect of the present invention is the posture control device according to any one of third to sixteen using the propulsive force amplifying device according to the seventeenth or eighteenth aspect as a wind flow generating device.

  According to this aspect, the propulsive force of the wind flow generator can be amplified.

Flying device of the 27th aspect of the present invention is a flying device according to any one of the 20-25 using the propulsion unit posture control apparatus according to the 26th.

  According to this aspect, the propulsive force of the propulsion unit can be amplified.

Flying device of the twenty-eighth aspect of the present invention is a flying device according to any one of the first 20, second 25 or third 27, further comprising an auxiliary blade so as counteract the moment about the overall center of gravity due to external wind flow wind pressure It is a thing.

  According to this aspect, the air pressure moment due to the external wind flow when the airframe starts the horizontal direction is canceled, and the airframe can perform a stable horizontal direction.

A propulsive force amplifying device according to a twenty- ninth aspect of the present invention is the propulsive force amplifying device according to the seventeenth or eighteenth aspect, wherein a moment around the total center of gravity of each propulsion unit due to external wind current is applied to each propulsion unit. Auxiliary wings are further provided so as to cancel.

  According to this aspect, the air pressure moment due to the external wind flow when the airframe starts the horizontal direction is canceled, and the airframe can perform a stable horizontal direction.

The posture control device according to the thirtieth aspect of the present invention is a nineteenth posture control device that satisfies the expression 21- (44) -3.

  According to this aspect, the posture control device can be stabilized against roll and counter-rotation due to propeller rotation.

The flying device according to the thirty-first aspect of the present invention is the twenty-fifth attitude control device, and satisfies Expression 21- (44) -3.

  According to this aspect, the posture control device can be stabilized against roll and counter-rotation due to propeller rotation.

The thirty-second aspect attitude control device of the present invention is a 30th posture control device, in parallel their center axes from each other, toward the lower side thereof the exhaust side with directing their intake side to the upper Each of the propulsion units spaced apart from each other, and a connecting member that interconnects the propulsion units, and each propulsion unit has a center of gravity and a wake wind pressure center point that coincide with each other. And is installed so that it can be rotated in the same direction around each wake wind pressure center point or each centroid at the same time, and the overall centroid and overall wake wind pressure center point of the attitude control device as a whole. And the general external wind pressure central point and the general external wind flow central point.

  According to this aspect, the attitude control device can be stably hovered and further stably shifted to horizontal flight.

The attitude control device of the thirty- third aspect of the present invention is the thirty- first flight device, with their central axes parallel to each other, their intake side facing upward and their exhaust side facing downward, Each of the propulsion units spaced apart from each other, and a connecting member that interconnects the propulsion units, and each propulsion unit has a center of gravity and a wake wind pressure center point that coincide with each other. It is set and installed so that each wake wind pressure center point or each center of gravity can be rotated in the same direction at the same time, and the overall center of gravity of the aircraft, the overall wake wind pressure center point, and the overall external wind pressure The center point and the central point of the general external wind flow pressure are made to coincide.

  According to this aspect, the flying device can be stably hovered and further stably transferred to horizontal flight.

It is a figure for demonstrating the propeller machine 40 which concerns on Embodiment 1. FIG. It is VI-VI sectional drawing of FIG. 1 and FIG. It is a figure explaining how to obtain | require the wind-pressure center point by the propeller wake. It is an example figure of a propeller. 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 α and the angle of blur at the tip of the propeller rotation shaft is β. It is the figure which showed the state in which the wind force with the force of mg acts on the stable blade 61. FIG. It is a perspective view of the airframe 170 at the time of arrange | positioning several n diagonal connection cylindrical propulsion units 170a-170d on a coaxial periphery. 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. FIG. 5 is a simplified front view showing a state in which all propeller rotation axes are inclined at an angle β in the same direction in an airframe 110B in which four cylindrical propulsion units 90a, 90b, 90a ′, 90b ′ are radially connected. It is a perspective view of airframe 110C which connected eight cylindrical propulsion units 90c-90j radially. FIG. 4 is a simplified front view of an airframe 110B in which four cylindrical propulsion units 90a, 90b, 90a ′, 90b ′ are radially connected. It is the simplified view of the side view which showed the flight state of the airframe 90 which consists of a single cylindrical propulsion unit. It is a perspective view of the airframe 130 which connected two cylindrical propulsion units 90a and 90b in parallel. It is a perspective view of the airframe 140 which connected three injection type propulsion units 140a-140c diagonally. It is a front view of the airframe 130B which diagonally connected two cylindrical propulsion units 90a and 90b. It is a top view in the horizontal flight of the airframe 150 which diagonally connected three cylindrical propulsion units 90a-90c. It is the schematic of the airframe 160 which connected the four cylindrical propulsion units 90a-90d diagonally. FIG. 6 is a plan view of the airframe 160 that is connected to the three cylindrical propulsion units 90a to 90c in an oblique manner and includes an auxiliary wing 160e during horizontal flight. It is the simplified view of the side view which showed the flight state of the body 180 which consists of a single cylindrical propulsion unit. It is the figure which looked at the airframe 170 of FIG. 10 from the top. It is a diagram of when the tubular propulsion unit 90 in FIG. 20 is rising at a upward velocity v _2. FIG. 6 is a simplified side view of an airframe 250 in which three, for example, cylindrical propulsion units (for example, 90) are connected in three oblique directions. It is an example figure of the flying car 300 using n diagonal connection same direction simultaneous rotation.

<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.

  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).

  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. .

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.

  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.

  Under the above setting, a condition for canceling the counter torque received by the propeller 43 by the propeller rotation is obtained.

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.

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. Then, the airflow density ρ ₀ is expressed by the equation (2).

Ltanα = r _0/2 So, put the tanα = ω, since the L = r ₀ / 2ω, equation (2) is the equation (3).

  Then, by substituting Equation (3) into Equation (1), the air flow density ρ becomes Equation (4).

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).

Since r _l = 2ωR, equation (6) is obtained.

  Therefore, Formula (7) is obtained.

  Substituting equation (7) into equation (5) yields equation (8).

Then, assuming that the total air volume directly under the propeller 43 is [ρ ₀ ], [ρ ₀ ] is expressed by Equation (9).

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).

  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.

  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.

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 ( The force F ′ is considered to be proportional to ρ _l (= ρ ₀ (r ₀ / r _l ) ² ) given by the equation (8) of the air flow density considering the wind spread.

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

Then, the moment M ₀ around the propeller rotation axis received by the propeller 43 is expressed by Equation (15).

Here, when considering the moment M _l around the propeller rotation axis on the horizontal line on the aircraft 41 at an arbitrary distance l (= nr ₀ ) below the propeller 43 when considering the aircraft 41 in FIG. 3. (16).

  It can be seen from Equation (15) and Equation (16) that Equation (16a) holds.

M ₀ = M _l (16a)
That is, from the equation (16a), it can be seen that the moment on the horizontal line at an arbitrary distance l on the main wing 41a matches the instantaneous moment of the propeller 43.

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.

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).

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.

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.

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).

  Equation (20) becomes Equation (45) when there are m main wings 41a.

  Next, based on FIGS. 2 and 3, the wind pressure center point for the propeller wake in the airframe 41 is calculated.

  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).

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.

  The moment Mac is expressed by equation (61).

  Similarly, the moment Mcb is expressed by Expression (62).

  And from Mac = Mcb, Formula (64) is materialized.

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.

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.

  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.

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.

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)). Therefore, the counter-rotation due to the reaction of the propeller rotation in the airframe 41 can be stopped simply by setting the height Δl of the vertical main wing 41a so as to satisfy Δl = 2πr ₀ / m.

  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.

  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.

<Embodiment 2>
As shown in FIG. 5, the propeller device 50B of this embodiment includes a plurality of layers of the airframe 41B of the propeller device 50B (n layers: n is an integer equal to or larger than 2; shown in FIG. 5 when n = 3) H1, H2 ,..., Hn.

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. The cylindrical body 47 _i is fixed to, for example, the side end surface of the partition plate 41 a _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 .

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.

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.

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.

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 .DELTA.l _1, .DELTA.l _2, ..., the number m _1, m ₂ of diameter units .DELTA.l _n and the vertical partition plate, ..., since m _n is set to satisfy equation (60), As in the case of the first embodiment, the counter torque due to the reaction of the propeller rotation 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.

<Embodiment 3>
In this embodiment, an attitude control device that is 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 S23.

S1. (Force F _C applied to wind pressure center point C by the propeller wake of the triangular wing aircraft)
Consider a triangular wing fuselage 63 as shown in FIG. The fuselage 63 has, for example, a propeller 60, a trapezoidal shape in a side view arranged on the lower side of the propeller 60, and a radial (for example, cross-shaped) lower stabilizer wing 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.

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: 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 wake wind pressure center point C n _W r ₀ : distance from the propeller 60 at the external wind pressure center point W

  Further, Formulas 1- (18) and 1- (1) are obtained from Embodiments 1 and 2.

  Experiments were performed with the airframe 63 of FIG. The numerical conditions of the airframe 63 at that time are as follows.

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).

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.

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.

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.

When the ratio between n _GC and n _GW when W is W = 1 is obtained, the coefficient W = 1.06 may be multiplied by n _GW in Equation 1- (2). As a result, Equation 1- ( 9) is obtained.

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.

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.

In this case, the relational expression between n _GC and n _GW is represented by Expression 1- (14).

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 1- (16). And Formula 1- (17).

  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. (Multiplier coefficient N when the number of diameter units of the lower stabilizer blade is m)
Here, let us consider a case in which six triangular wings (upper stabilizing wing 62 and lower stabilizing wing 61) are radially provided in the fuselage 63 of FIG. 6 as shown in FIG. The wind pressures caused by the propeller wakes applied to the stabilizing blades 61-1, 61-2, 61-3, 61-4, 61-5 and 61-6 of the lower stabilizing blade 61 are respectively expressed as F _C-1 and F _C-2. , F _C-3 , F _C-4 , F _{C-5 and FC-6} . In addition to the wind pressure F _C-1 due to the wake of the propeller applied to the stabilizer blade 61-1 itself, the stabilizer blade 61-1 has the stability at the wind pressures FC _-2 to FC _-6 related to the stabilizer blades 61-2 to 61-6. A component perpendicular to the wing 61-1 is applied. In consideration of this, the total force [F _C ] of the wind pressure applied to the stable blade 61-1 is expressed by Equation 2- (1).

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). .

  Here, if N is placed in {} of Equation 2- (2), Equation 2- (2) becomes Equation 2- (3).

  In general, it can be seen that N is expressed by Equation 2- (4), where m is the number of stabilizing blade diameter units of the lower stabilizing blade (radial stabilizing blade) 61.

  In the case of FIG. 7, since the relational expression 2- (8) is established, N in that case is N = 2.

  In general, 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. (Center point of wind pressure due to the wake of the propeller in the radial stable blade)
As described above, the distance from the propeller 60 to the wind pressure center point C due to the wake of the propeller is expressed by Expression 1- (1).

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 materialized 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).

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.

However, if the fuselage 63 is formed in a cylindrical shape so that the propeller wake does not diffuse, the propeller wake becomes a uniform parallel flow on the inner side of the cylinder as will be demonstrated in later experiments, and the wind pressure caused by 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. (Correction of F _C definition)
In the hovering experiment of the fuselage 63 in FIG. 6 in the absence of wind, the wind pressure F _C caused by the propeller wake is applied to the wind pressure center point C caused by the propeller wake in the lower stabilizing blade 61 and the center of gravity G of the fuselage 63 is expressed by Formula 1− ( As long as 17) is not satisfied, the following 1-4 can be estimated from the fact that the airframe 63 becomes unstable.

1. 1. The reason that the airframe 63 is unstable is the wake of the propeller. 2. The force of F _C can be assumed to be a force (pseudo lift) similar to lift because the wake behind the propeller flows almost parallel to the lower stabilizer blade 61. The fact that the pseudo-lift force is applied to the lower stabilizing blade 61 in the parallel flow means that the wake behind the propeller considered to be a parallel flow is actually not parallel to the lower stabilizing blade 61 but at an angle. What will happen 4. The reason why the propeller wake has an angle with respect to the lower stabilizer blade 61 is that the rotating shaft of the propeller 60 was originally inclined at a certain angle with respect to the lower stabilizer blade 61 or the tip of the rotating shaft due to the rotation of the propeller 60. Be considered to be blurring.

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.

The defining formula of F _C is Formula 3- (1).

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.

  Subsequent experiments prove that equations 4- (2) and 4- (3) are nearly correct. Furthermore, from these experiments, it is proved that the pseudo lift coefficient k is k≈1.

S5. (Determination of pseudo lift coefficient k)
FIG. 8 is a schematic side view of the airframe 63 when the inclination of the propeller rotation shaft, which is an unstable element during hovering of the airframe 63, is α, and the angle of blurring of the tip of the propeller rotation shaft during propeller rotation is β. is there.

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). Here, when α is an extremely small value, Expression 5- (9) becomes Expression 5- (10).

F _C is given by the equation 4- (2) as described above. The F _P is given by the formula 5- (11) from the balance between the vertical component and the weight mg of the fuselage 63 of the thrust F _P.

  Considering these things, Equation 5- (9) becomes Equation 5- (1).

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.

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).

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 stabilizer 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.

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.

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 airframe 63 becoming unstable, 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.

From the above results, the relational expressions of F _C and F _W are summarized as follows.

  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) 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. (Identity of the pseudo-lift F _C)
In order to confirm whether or not the pseudo lift F _C is really a commonly known lift, it was confirmed by an experiment whether the lift applied to the flat plate is applied to the lower stabilizing blade 61. As a result, it was found that no force such as lift was applied to the propeller 60 at any angle with respect to the stable blade 61.

As I have felt several times in the experiments so far, if I try to translate the aircraft 63 with a 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.

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 direction of 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.

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.

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.

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 formula 6- (2) is an equation of n _GC in this experiment calculated from equation 5 (5), the point W is arranged at a distance n _GW r ₀ determined by the n _GW When the hovering experiment was performed again, the fuselage 63 hovered stably.

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.

Here if F _C becomes the force is resistive force applied to the body 63 when the body 63 begins to move, it is necessary to correct the basic formula 5- (6) of the F _C. The formula 5- (6) is corrected using FIG.

Sign a in FIG. 9 (= gtanβ), when the angle of deflection of the propeller shaft 60a distal is beta, represents the acceleration body 63 begins to move in the horizontal direction by the propeller thrust F _P is inclined by an angle beta.

Deflection of the propeller shaft 60a is not rotated in all directions, the vertical component of the propulsion force F _P is supporting the weight of the aircraft 63. Therefore, the wind force flowing downward from the propeller 60 flows along the stable wing 61 of the airframe 63 with an average force of mg on the lower side. At this time, the airframe 63 starts to move in the horizontal direction at the acceleration a given by Equation 6- (4).

Therefore, the wind having the total power mg flowing downwards hits the stable wing 61 of the body 63 at an angle β as shown in FIG. From this fact, the basic formula of F _C in Expression 5- (6) is corrected as shown in Expression 6- (5) or Expression 6- (6).

The above results show important facts. That is, considering the case where the tip of the propeller rotating shaft 60a is tilted by an angle β only in one direction instead of front, rear, left and right, and when the fuselage 63 is flying horizontally, F in Equation 6- (6) _P is the propulsive force of the airplane 63 in the horizontal direction, and the angle β is considered to be the angle at which the main wing 61 of the airplane 63 is inclined toward the traveling direction, that is, the angle of attack of the main wing 61. -(6) is considered to be an expression of lift.

In general, airplanes with fixed wings gain lift by moving horizontally. In such an airplane, the fixed wing moves horizontally in the air with an angle of attack, and the fuselage 63 in which the stable wing 61 is disposed along the propeller rotation shaft 60a has a propeller rotation shaft during hovering. The state of starting to move in the horizontal direction due to the shake of 60a is considered to be exactly the same with respect to the force applied to each of the fixed wing of the airplane and the stable wing 61 of the fuselage 63. From this, the expression 6- (6) is a propulsive force that is a completely different factor from the generally-described lift force (in other words, the pressure of the wind flowing from the circle having the diameter r ₀ (= the diameter of the propeller 60)) It can be said that this is an expression expressed by a multiple coefficient of the diameter r ₀ of the propeller 60.

  From the above results, the moment balance equation of Equation 5- (10) is similarly Equation 6- (7).

  When Expression 6- (15) and Expression 6- (16) are substituted into Expression 6- (7), Expression 6- (2) and Expression 6- (3) are replaced with Expression 6- (8), Expression 6 -(9).

Further, the basic formula of _FW is represented by formula 6- (10).

When there is no π-fold effect, the basic expressions of F _C and F _W are Expressions 6- (11) and 6- (12), respectively.

  If the above theory is correct, in the above experiment, when the center of gravity G is located above the wake wind pressure center point C, the external wind wind pressure center point W was at the bottom, so the moment balance In addition, the shaking of the airframe 63 becomes larger and the airframe 63 should become unstable. However, as a result, the fuselage 63 was only temporarily, but was stable.

Therefore, in both cases where the center of gravity G is disposed above the wake wind pressure center point C and where the center of gravity G is disposed below the wake wind pressure center point C, the n _GC in each case The experiment was performed again by placing the external wind wind pressure center point W below the wake wind pressure center point C at the corresponding n _GW distance. As a result, when the center of gravity G is placed below the wake wind pressure center point C, the fuselage 63 performs stable hovering for a long time, but when the centroid G is placed above the wake wind pressure center point C. The fuselage 63 did a stable hover for a while, but suddenly started to sway from side to side.

  What can be considered from this result is that the aircraft 63 should first start to rotate around the horizontal line passing through the center of gravity G immediately before the aircraft 63 starts to move in the horizontal direction due to the swing of the tip of the propeller rotary shaft 60a. This is probably because a resistance force is generated at the wake wind pressure center point C and the external wind wind pressure center point W so that a moment opposite to the rotation is generated. Therefore, when the center of gravity G is located above the wake wind pressure center point C, the fuselage 63 performs stable hovering for a while, but it balances with some trigger (for example, a slightly strong flow of ambient air). If broken, it will no longer be possible to return to the original stable state.

Due to this fact, the propeller wake gives a force expressed by Equation 6- (13) to the stable blade 61 when a force of F _V acts on the stable blade 61 from the outside perpendicularly, On the other hand, when there is no π-fold effect, it can be seen that the force of Expression 6- (14) is applied to the stable blade 61.

The addition the above facts, even if the machine body 63 but of course if you are moved in the horizontal direction F _V becomes a force, which is rotated about the center of gravity G added is F _V becomes a force to the tip of the body 63 It can also be seen that the moment of the fuselage 63 can be calculated assuming that the wake wind pressure center point C is subjected to the resistance force F _C represented by the equation 6- (13) or the equation 6- (14).

S7. (Cylindrical stabilizer)
Consider an aircraft 70 as shown in FIG. The airframe 70 includes a propeller 71, a rectangular shape in a side view disposed below the propeller 71, for example, a cross-shaped lower radial stabilizer wing 72, the same height as the lower radial stabilizer wing 72, and the lower radial stabilizer For example, a cylindrical lower cylindrical stabilizing blade 73 disposed coaxially so as to surround the periphery of the blade 72 and a coaxial line disposed above the propeller 71 and having the same diameter as the lower cylindrical stabilizing blade 73. For example, a cylindrical upper cylindrical stabilizing blade 74, a rod-shaped connecting member 76 that connects the respective 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.

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.

  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 condition values of the experiment are n = π, N = 1, n _X = 0.7134, r ₀ = 15.24 cm.

Also, here, the following two assumptions 1 and 2 are set, the values of n _GC and n _GW are obtained using Equations 6- (8) 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.

  The above two assumptions 1 and 2 are as follows.

Assumption 1: Wind pressure F _C due to the propeller wake applied to the lower cylindrical stabilizer wing 73 is applied to one of the lower radial stabilizer wings 72 (that is, the height of the lower cylindrical stabilizer 73 × the area of the diameter). Same as F _C.

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 73 stable blade is also added. In other words, an area that is twice the area of one stabilizing blade of the lower radial stabilizing blade 72 is S _C.

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).

  When the fuselage 70 was adjusted to satisfy the above condition values and calculated values, the fuselage 70 was stably hovered.

  In the experiment of the airframe 70 in FIG. 11, the hovering was stably performed without any left / right shaking, but the airframe 70 was rotating 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.

  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.

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.

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, an area that is three times the area of one stabilizing blade of the lower radial stabilizing blade 72 is S _C.

  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.

The diameter of each of the lower cylindrical stabilizer wing 73 and the upper cylindrical stabilizer wing 74 and the lateral width of the lower radial stabilizer wing 72 are both increased to 1.115r ₀ (r ₀ = 15.24 cm), and the first experiment and 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.

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.

Further, when an in-cylinder cylindrical stabilizer blade (diameter R ₁ , height h ₁ ) is combined in the lower cylindrical stabilizer blade 73, Equations 7- (4) to 7- (6) are obtained.

  Further, when the lower cylindrical stabilizer 73, the in-cylinder stabilizer and the radial stabilizer 72 are combined, Expressions 7- (7) to 7- (9) are obtained.

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.

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.

Here, if the expression 7- (13) is put, S _C becomes the expression 7- (14), and if the expression 7- (15) is put, the expressions 7- (16) to 7- (18) are obtained. Expressions 6- (2) and 6- (3) become Expressions 7- (19) and 7- (20).

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. (Anti-torque canceling condition due to propeller rotation of an aircraft equipped with cylindrical stabilizer blades and radial stabilizer blades)
The airframe 70 in FIG. 11 performed hovering stably, but rotated forward in the same direction as the propeller rotation. The n value of the height of the lower radial stabilizer blade 72 for stopping this forward rotation is obtained.

  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).

  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.

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.

“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.

  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.

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. (Correction of F _C = HπF _P sinβ)
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?

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.

  When β in Expression 9- (1) is calculated from the experimental value of S8, β≈10.95 °.

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.

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 next, 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 °.

  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.

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).

  From this result, it can be said that the pitch angle 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.

  From this result, it can be seen that it is almost close to an 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).

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).

  If the number of the lower radial stabilizer blades is m, Equation 9- (9) is obtained.

  In the case of a propeller with a pitch of 5 inches, the n-value of the lower radial stabilizer blade for stopping counter-rotation has a width, and when the n-value is between about 1.4 and 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.

  Here, the difference in physical phenomena between Equation 6- (6) and Equation 9- (1) will be considered. Equation 6- (6) can be rewritten as Equation 9- (10).

  The difference between Equation 9- (1) and Equation 9- (10) is the presence or absence of the element (1-sinβ). So what does (1-sinβ) mean? In the case of Expression 9- (10), even if the tip of the propeller rotation shaft is swung by the angle β, the wind emitted from the propeller is blown down at an angle of 0 degrees on average. Equation 9- (10) is a wind pressure equation received from the wind flowing on the stable blade at an angle of 0 degrees.

  On the other hand, Formula 9- (1) is a wind pressure formula when the wind emitted from the propeller hits the stable blade at the propeller execution angle β. Considering these things, when the wind is flowing at the speed v, the angle β is the wind pressure when the force of the wind hits the stable blade at the angle β, as shown in Equation 9- (1) It can also be said that the formula 9- (10) is satisfied.

  In the case of Equation 9- (1), it can be interpreted that the original wind hits the stable blade at an angle β, and the component perpendicular to the stable blade of the force that the wind has exerts a force on the stable blade. In the case of Expression 9- (10), as is clear from FIG. 9, the stabilizer impinges toward the wind flowing in parallel with itself due to the swing of the tip of the propeller rotating shaft. As a result, it can be interpreted that the tan β component of the force contained in the wind exerts a force on the stable blade.

  Therefore, the element of (1-sinβ) is an element that appears by the angle of the wind flow direction with respect to the stable blade. In the case of Expression 9- (10), the direction of the wind flow is very close to 0 ° with respect to the stable blade. Therefore, (1-sin β) ≈1, and it is considered that the element of (1-sin β) did not appear. Then, when the wind current hits the stable wing at an angle β, why does the element (1-sinβ) appear?

Now, an expression expressing the force F _P of the wind flow having the wind speed v _P by v _P is Expression 9- (11).

  When Expression 9- (11) is substituted into Expression 9- (1), Expression 9- (12) is obtained, and the first term on the right side is the same as 9th- (10).

However, if you look at the second term on the right side of Equation 9- (12), the meaning of this term is Hπ times the force that the wind current of the component perpendicular to the stable blade of the wind flow with the wind speed v _P has. I understand. Therefore, the meaning of Equation 9- (10) is that when the wind flow with the wind velocity v _P hits the stable blade at an angle β, the component perpendicular to the stable blade of the wind flow bounced off from the stable blade has the force that the original wind flow has. It can be interpreted that the component perpendicular to the stable wing is reduced.

S12. (Physical meaning of F _C = nπmgtanβ)
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.

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.

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.

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).

The formulas for _FW and W are also given by formulas 12- (3) to 12- (5).

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 as in Expressions 12- (6) to 12- (9).

Previous theories from above, it can be seen hovering by wind ejected from the hole of diameter r ₀ not only propeller motor (e.g. a rocket, hovering machine according jets and gas injection, etc.) also are applicable.

S13. (When the lower cylindrical stabilizer is extended above the propeller)
The airframe 80 in FIG. 12 has a propeller 81, a rectangular shape in a side view disposed below the propeller 81, for example, a cross-shaped lower radial stabilizer wing 82, and a coaxial line so as to surround the periphery of the lower radial stabilizer wing 82 A cylindrical stabilizer wing 83 whose lower end is the same height as the lower radial stabilizer wing 82 and whose upper end is extended above the propeller 81, and a propeller disposed on the lower radial stabilizer wing 82. Drive unit (not shown).

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.

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 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).

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.

  As expected, the aircraft 80 hovered very stably and no anti-rotation occurred as expected. 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- (20). However, in the experiment without wind, it seems that the most stable hovering has been performed so far. Looked.

  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, stable hovering was performed 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.

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. (Attitude control device by combining various types of stabilizing blades)
The fuselage (attitude control device) 90 of FIG. 13 is a cylindrical stabilizer wing 91 and one coaxially arranged along the central axis 92 of the cylindrical stabilizer wing 91 inside the cylindrical stabilizer wing 91. The radial stabilizers 93 and 94 that are symmetrical with respect to the central axis as described above (two in this example), for example, are rectangular in side view, and one or more (two in this case) that are coaxially disposed inside the cylindrical stabilizer 91. In-cylinder cylindrical stabilizer blades 95, 96, a propeller 97 disposed on the central axis of the cylindrical stabilizer blade (eg 91), and a propeller drive unit disposed on the radial stabilizer blade (eg 93) (Not shown). Here, each of the cylindrical stabilizing blades 91, 95, 96 is formed in a cylindrical shape, for example. The diameter of the propeller 97 is assumed to be smaller than the diameter of the cylindrical stabilizing blade 91.

  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.

In this airframe 90, the wind pressure center points H ₀ and C ₁ caused by the propeller wakes in each of 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 rotation axis 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.

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) Increase or decrease the number of stabilizing blades of the radial stabilizing blades 93 and 94 (2) Increase or decrease the height of the radial stabilizing blades 93 and 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).

  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 stable blade 91.

  The airframe 90 configured in this way performs hovering with stability, has resistance against the influence of external wind, and can perform hovering while maintaining surprising stability.

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.

  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.

  Further, since one or more in-cylinder cylindrical stabilizer blades 95 and 96 are coaxially provided inside the cylindrical stabilizer blade 91, the stability of the airframe 90 can be further improved in the wind flow.

  Further, since the propeller 97 is provided, 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 | positions so that (5) may be satisfy | filled, the rolling of the body 90 can be prevented.

  Further, since it is designed so as to satisfy Expression 14- (1), the counter-torque caused by the propeller rotation can be offset, and the aircraft body 90 can be prevented from counter-rotating due to the propeller rotation.

S15. (Usage control device usage example)
The airframe (attitude control device) 110 in FIG. 14 is a combination of two airframes 90 of S14 (hereinafter referred to as airframes 90a and 90b). More specifically, the airframe 110 has its intake-side open end directed upward, its exhaust-side open end directed downward, and its intake-side open ends inclined in opposite directions (inclination angle 0). In addition, the two airframes 90a and 90b arranged at a distance from each other and a connecting member 111 that interconnects the airframes 90a and 90b are provided.

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.

  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.

Note that points P ₁ and P ₂ in FIG. 14 are the center points (propulsion points) of the propeller 97 of the airframes 90a and 90b, respectively, and α is the line segment G ₁ G ₀ (and the line segment G ₂ G). ₀ ) and the horizontal direction, λ 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. The angle between ₀ and the line segment P ₁ G ₀ and the angle between the line segment G ₂ G ₀ and the line segment P ₂ G ₀ .

Here, the line segment G ₀ P ₁ = the line segment G ₀ P ₂ = L, and the propulsive forces of the airframes 90a and 90b are F _P1 and F _P2 (where F _P1 = F _P2 ), respectively. Consider a state in which the angle γ is inclined toward one of the airframes (for example, 90a).

In the airframe 110 of FIG. 14, the moments around the center of gravity of F _P1 and F _P2 are always constant regardless of the value of γ, but the gravity applied to each of the center of gravity G ₁ , G ₂ and the driving force F _{When the} restoring moment ΔM due to the couple between _P1 and F _P2 is calculated, Equation 15- (1) is obtained and is not constant with respect to the value of γ.

_{Assuming that} the resultant force [F] of the propulsive forces F _P1 and F _P2 is balanced with the gravity mg of the airframe 100, Expressions 15- (2) and 15- (3) are established.

  Since θ = π / 2− (α + λ), Expression 15- (3) becomes Expression 15- (4).

  Each term in parentheses in the expression 15- (4) becomes the expression 15- (5), so the expression 15- (4) becomes the expression 15- (6).

From Equation 15- (6), the restoring force against the left and right shaking of the airframe 110 is that the resultant force of the propulsive forces F _P1 and F _P2 of the two cylindrical propulsion units 90a and 90b is as if from the center of gravity G ₀ (line segment P ₀ It can be seen that it is given as acting on a point (action point P) positioned upward by a distance of G ₀ + line segment O ₀ G ₀ ).

Next, as shown in FIG. 19, two sets of airframes 110 (which are referred to as airframes 110 and 110 ′) are connected in a cross shape in plan view with their center of gravity G ₀ matched (hereinafter referred to as a cross-shaped connection). 110B). In FIG. 19, the cylindrical propulsion units 90a, 90b, 90a ′, 90b ′ of the airframes 110, 110 ′ are shown as straight lines for convenience. In the cross-shaped connecting unit 110B, a restoring moment [ΔM] due to gravity is calculated. When viewed from the front (cylindrical propulsion unit 90a ′ side), when the cross-shaped connecting unit 110B is tilted to the left and right by an angle γ, the restoring moment ΔM ₁ by the left and right tubular propulsion units 90a and 90b is expressed by the equation 15− ( Since it is obtained by substituting mg of 4) with mg / 2, the formula 15- (7) is obtained.

At this time, the restoring moment ΔM ₂ by the two front and rear cylindrical propulsion units 90a and 90b when the cross-shaped connecting unit 110B is viewed from the front is expressed by Expression 15- (8).

  Since α + θ = π / 2−λ, Expression 15- (8) becomes Expression 15- (9).

  Therefore, the total restoring moment [ΔM] of the cross-shaped connecting unit 110B is expressed by Expression 15- (11).

From Expression 15- (11), the restoring force against the left and right shaking of the cross-shaped connecting unit 110B is the sum of the thrusts of the four cylindrical propulsion units 90a, 90b, 90a ′, 90b ′ as if the center of gravity G ₀ ( It can be seen that it is given as acting on a point located upward by a distance of line segment P ₀ G ₀ + 1/2 · line segment O ₀ G ₀ ).

Here, when the airframe (airframe combining two cylindrical propulsion units 90a and 90b) 110 in FIG. 14 is tilted by an angle γ in the opposing direction of each cylindrical propulsion unit 90a and 90b, each cylindrical propulsion unit 90a, If the F _C becomes the force applied to 90b and F _C1, F _C2, consider what their resultant force F _C0 of each force F _C1, F _C2 becomes what kind.

Assuming that the propulsive forces of the cylindrical propulsion units 90a and 90b are F _P1 and F _P2 and the propulsive forces F _P1 and F _P2 are balanced with the gravity mg of the fuselage 110, the relationship of Equation 15- (12) is established. obtain.

The respective tubular propulsion unit 90a, F _C becomes the resistance force F _C1 applied to the flow wind pressure center point C _1, C ₂ after 90b, F _C2 is given by the formula 15 (34).

Further, the resultant force (total wake wind pressure) F _C0 of F _C1 and F _C2 is applied to the point C ₀ in FIG. 14 and is therefore expressed by Equation 15- (13). Therefore, the total external wind pressure F _W0 is expressed by Equation 15- (14). Become. The point C ₀ is the total wind pressure center point when the airframe 110 is viewed from the front (that is, when viewed from the cylindrical propulsion unit 90a ′ side), and passes through the point C ₁ and the center of the cylindrical propulsion unit 90a. It is given by the intersection of a straight line perpendicular to the axis 92 and a straight line passing through the point C2 and perpendicular to the central axis 92 of the cylindrical propulsion unit 90b.

This airframe 110 can be regarded as one propulsion unit in which a force of F _C0 represented by Expression 15- (13) is applied to the point C ₀ with respect to left and right shaking.

However, forces acting the body 110 Overall wind pressure center point when viewed from the side (i.e. when viewed from the cylindrical propulsion unit 90a or 90b side), C ₁₂ next point in FIG. 14, in the overall wind pressure center point Is the same as the force expressed by Equation 15- (13). The point C ₁₂ is an intermediate point between the points C ₁ and C ₂ . Therefore, the fuselage 110 is unstable because the position of the total wind pressure center point changes depending on the viewing direction.

In order to eliminate this instability, when the four cylindrical propulsion units 90a, 90b, 90a ′ and 90b ′ are combined as in the cross-shaped connecting unit 110B, the total wake wind pressure center point of the four cylindrical propulsion units is combined. C appears at the midpoint of the line segment C ₀ C _{12 in} FIG. The position of this point C does not change in any direction of the cross-shaped connecting unit 110B, and is always located at the midpoint of the line segment C ₀ C ₁₂ .

For this reason, when combining a plurality of cylindrical propulsion units, it is desirable that the cross-shaped connecting unit 110B is a single unit (combination A). When only two cylindrical propulsion units are combined as in the fuselage 110 of FIG. 14, the inclination angle α of these two cylindrical propulsion units is set to 0 °, and the total wind pressure center of these two cylindrical propulsion units is set. When the points C ₀ and C ₁₂ are made to coincide with each other, the position of the total wind pressure center point C ₀ does not change when viewed from any direction of the entire fuselage 110, so that the fuselage 110 can be stably hovered (combination) B). In both combinations A and B, the center of gravity G ₀ and the external wind pressure center point W ₀ of the entire airframe must be arranged so as to satisfy the following equation 19- (5).

  FIG. 17 is a simple diagram in which each cylindrical propulsion unit 90a, 90b, 90a ′, 90b ′ is simplified and represented by a straight line when the cross-shaped coupling unit 110B is viewed from the front (cylindrical propulsion unit 90a ′ side). is there. The central cylindrical propulsion unit is represented by a single line because the front and rear cylindrical propulsion units 90a 'and 90b' appear to overlap each other.

Now, it is assumed that the swing angles of the propeller rotary shaft tips of all the cylindrical propulsion units 90a, 90b, 90a ′, 90b ′ are swung, for example, by β on the left side of FIG. 7 at the same timing. General After Luke pressure at this time (i.e. the tubular propulsion units 90a, 90b, 90a ', 90b ' Luke total force of the pressure after applied to) F _C is whether appear in any position in any size Think.

Each tubular propulsion units 90a, 90b, 90a ', 90b ' flow wind pressure center point C _{1 after, C 2, C 3, C} 4 to take F _C becomes a force (back Luke pressure) to F _C1, F _C2, F _{When C3} and F _C4 are set, the magnitude of the two resultant forces of F _C1 and F _C2 is F _C0, and the magnitude of the two resultant forces of F _C3 and F _C4 is F _C34 , F _C0 is expressed by Equation 15- (15 ) However, F _P1 = F _P2 = F _P3 = F _P4 = (mg / 4) · 1 / (cosα · cosβ).

F _C34 is expressed by Expression 15- (16).

The relation of the formula 15 (15) (16) from Equation 15 (17) is satisfied, overall after Luke pressure F _C becomes Equation 15 (18).

From Expression 15- (17), it can be seen that the action point of F _C is the midpoint between the point C ₀ and the point C ₃₄ .

Further, Formula 15- (18) indicates that when the deflection angles of the propeller rotary shaft tips of the respective cylindrical propulsion units 90a, 90b, 90a ′, 90b ′ are all β (note that these cylindrical propulsion units have the same timing) In the case of shaking in the same direction, the instability is maximized. Therefore, in the case of shaking at the same timing, and for the sake of simplification, all the swing angles are considered to be the same β.), F _{C of} Equation 15- (18) , it is seen that Luke pressure F _C and direction nor magnitude same after the time of the deflection angle of the propeller shaft leading end of the propulsion unit when considering the entire body 110B as a single propulsion unit beta. However, the difference is that the restoring force by gravity against the shaking of the airframe 110B becomes very large.

There are two important things here. First, in the case of a fuselage in which a plurality of cylindrical propulsion units are connected radially in plan view, such as the cross-shaped coupling unit 110B, these cylindrical propulsion units are in the horizontal direction and the vertical direction of the entire fuselage. When the counter-rotation around the central axis starts, a resistance force of F _C is generated and the resistance force becomes a very large resistance moment, and therefore the anti-torque canceling force by the radial stabilizer blades in each cylindrical propulsion unit In combination, the anti-rotation tendency of the entire aircraft is almost eliminated.

Secondly, in the case of the fuselage 110 (FIG. 14) in which two cylindrical propulsion units 90a and 90b are connected, the position of the total wake wind pressure center point is point C _{0 in} FIG. since body 110 becomes unstable changes between bets point C _12, to eliminate the instability, another set of tubular propulsion unit 90a as the cross-shaped connection unit 110B of FIG. 17 ', 90b 'Is combined in a cross shape in plan view, and the total wake wind pressure center point is fixed at point C.

However, the two sets of cylindrical propulsion units 90a, 90b and 90a ′, 90b ′ are completely separate, and in fact, these two sets of cylindrical propulsion units are each provided with their own total wake wind pressure center point (hereinafter referred to as the center point). , The total wake wind pressure center point of each set unit is referred to as the set unit total wake wind pressure center point). For example, the left and right swings when viewed from the front as shown in FIG. The group-wise integrated wake wind pressure center point of the two cylindrical propulsion units 90a, 90b is a point C ₀ , and the group-unit integrated wake wind pressure center point of the two cylindrical propulsion units 90a ′, 90b ′ is a point C ₃₄ This means that the forces F _C0 and F _C34 of F _C applied to the two set unit total wake wind pressure center points C _{0 and} C ₃₄ appear at the point C in appearance.

The above point C ₀ is a straight line passing through the wake wind pressure center point C ₁ of the tubular propulsion unit 90a and perpendicular to the central axis 92 of the tubular propulsion unit 90a, and the wake wind pressure center point of the tubular propulsion unit 90b. The point C ₃₄ is an intersection with a straight line passing through C ₂ and perpendicular to the central axis 92 of the cylindrical propulsion unit 90b, and the above point C ₃₄ is the wake wind pressure center point C ₃ , C of each of the cylindrical propulsion units 90a ′, 90b ′. The midpoint of ₄ .

Here, when the cross-shaped connecting unit 110B is rotated 45 ° horizontally (however, the inclination β of the rotation axis is always inclined by β as shown in FIG. 17 when viewed from the direction of rotation). The above-mentioned two sets of unit wake air pressure center points C _{0 and} C ₃₄ gradually approach each other with respect to the left and right sway of the aircraft as viewed from that direction, and viewed from the direction of 45 ° rotation. When the left and right swings at that time, the above-mentioned two sets of unit wake total wind pressure center points C _{0 and} C ₃₄ overlap with the point C. When this is further rotated by 45 °, one set unit overall wake wind pressure center point that has been lowered further falls further down, and the other set unit overall wake wind pressure center point that has been raised is further raised. The center wake wind pressure center points C _{0 and} C ₃₄ for the left and right swaying when viewed from the direction fully turned 90 ° are respectively reversed from the initial positions. Become.

With respect to the left and right shaking of the airframe when FIG. 17 is viewed from the front or the side, the above-described two unit-unit total wake wind pressure center points C _{0 and} C ₃₄ are farthest apart, and the airframe 110B is located at the midpoint. Even if the entire center of gravity G is arranged and the external wind pressure center point W of the entire body 110B is made to coincide with the center of gravity G, the two sets of unit-wise total wake wind pressure center points C ₀ that are separated from each other with respect to the swing of the body 110B. , from the viewpoint of stability against shaking of the made (i.e. aircraft that is balanced by F _C becomes the force F _C0, F _C34 applied to C _34, each set unit overall wind pressure center point is important that you are away each other is there). This leads to a reduction in size and weight of the fuselage 110B.

  Further, by replacing each propeller type cylindrical propulsion unit 90a, 90b, 90a ′, 90b ′, which is a wind flow generator, with an injection type propulsion unit (such as a gas injector or a rocket injector), even in a vacuum, that is, in space. It becomes available.

However, when the fuselage 110B of FIG. 17 is rotated by 45 ° in the horizontal direction, the above two unit-unit total wake wind pressure center points C ₀ and C ₃₄ overlap with the midpoint C thereof. It becomes very unstable with respect to the left and right shaking seen from the 45 ° direction. In order to eliminate this instability, another set of cross-shaped connecting units 110B are concentrically connected, that is, a total of eight cylindrical propulsion units 90 (90c to 90j) are radially formed as shown in FIG. What is necessary is just to comprise the radial 8 connection unit 110C by connecting with the connection member 110a. This is because when eight cylindrical propulsion units 90c to 90j are connected radially, there are at least two group unit rear wind pressure center points in any direction, and after four unit units are combined in the other directions. This is because there is a flow wind pressure center point, and a resistance force of F _C is generated in a balanced manner at each set unit overall wake wind pressure center point.

  In FIG. 18, the connecting member 110 a is configured by combining four trapezoidal plate members 110 b to 110 e in a radial manner with their center axes coinciding with each other. And the cylindrical propulsion units 90c-90j are arrange | positioned by the hypotenuse of the both sides of each board member 110b-110e.

  From the above, in order to construct a stable airframe by combining four or more cylindrical propulsion units, an even number of cylindrical propulsion units (preferably eight or more even number of cylindrical propulsion units) are connected radially. Further, the center of gravity of the entire aircraft is shown in FIG. 17 so that straight lines drawn perpendicularly to the central axis (cylinder axis) of the cylindrical propulsion unit from the wake wind pressure central point of each cylindrical propulsion unit are concentrated at one point. The upper and lower set unit total wake wind pressure center point may be arranged as shown in FIG. By doing so, as will be understood in later experiments, it is not necessary to match the center of the external wind pressure of the entire aircraft with the center of gravity of the entire aircraft. Can stabilize the aircraft.

FIG. 19 is a simple view of the same cross-shaped connecting unit 110B as FIG. 17 as seen from the front. Since the cylindrical propulsion units 90a ′ and 90b ′ appear to overlap, they are represented by a single straight line. The propulsive force of each cylindrical propulsion unit 90a, 90b, 90a ′, 90b ′ is set to F _P1 , F _P2 , F _P3 , F _P4 = F _P. Point C ₃₄ is the total wind pressure center point (set unit total wind pressure center point) of each cylindrical propulsion unit 90a ′, 90b ′ as viewed from the front direction, and point C ₀ is the same as each cylindrical propulsion unit 90a, 90b is the total wind pressure center point (set unit total wind pressure center point). F _Cb is a force of F _C applied to the point C ₃₄ when the cross-shaped connecting unit 110B starts to swing counterclockwise due to Δf _P , and F _Cf is a force of F _C also applied to the point C _0. . The distance from the center of gravity G ₀ to the propulsive force acting point (propeller center point) P ₁ or point P ₂ is L. The total center of gravity G ₀ of the airframe 110B is at the midpoint of the line segment C ₀ C ₃₄ .

From the state machine body 110B is performing a stable hover as shown in FIG. 19, to increase driving force of the tubular propulsion unit 90b from F _P only Delta] f _P, aircraft 110B consider a state swaying left and right.

The force of F _C due to the rotational moment due to Δf _P is divided into two, F _Cb and F _Cf in FIG. Angular acceleration when the by Delta] f _P body 110B is rotated, since the same in terms of aircraft 110B throat, if magnitude of the rotational moment at the point P ₂ is known, the propeller working point P _1, P _2, P ₃ and ₄ can also calculate the force that is assumed to be virtually applied. The virtual force applied to the cylindrical propulsion unit 90a ′ or 90b ′ is now Δf _34, and between the total action point P of each propeller action point P ₁ , P ₂ , P ₃ , P ₄ and the total center of gravity G _0. If the distance is n _G , the rotational moment ΔM _P2 due to Δf _{P at the} point P ₂ is expressed by Equation 15- (19).

Further, the rotational moment ΔM _{34 at the} point P ₃ or the point P ₄ is expressed by Expression 15- (20).

Since ΔM _P2 = ΔM ₃₄ must be satisfied, Expression 15- (21) is obtained.

Since Δf _{34 in} Expression 15- (21) is applied to both points P ₃ and P ₄ , F _Cb is expressed as Expression 15- (22).

Similarly, when virtual forces applied to the propeller action points P ₁ and P ₂ of the cylindrical propulsion units 90a and 90b are Δf ₁ and Δf ₂ , Expression 15- (23) is obtained.

Therefore, F _Cf is expressed by Equation 15- (24).

The couple moment ΔM of the two F _Cb and F _Cf is expressed by Expression 15- (25), where the length of the line segment C ₀ C ₃₄ is n _C34 .

n _C34 is obtained by drawing from FIG.

  Therefore, Expression 15- (27) is obtained from Expression 15- (22), (24), (25), (26).

In order for the airframe 110B to perform stable hovering, ΔM in Expression 15- (27) only needs to be larger than ΔM _P2 , that is, Expression 15- (28) is satisfied. Further, the relationship of Equation 15- (29) is obtained from Equation 15- (28).

Lcos θ is the distance from the center of gravity G ₀ to the cylindrical axes of the cylindrical propulsion units 90a, 90b, 90a ′, 90b ′. When Lcosθ = L _G, to obtain formula 15 and (30) from equation 15- (29).

From Equation 15 (30), each tubular propulsion units 90a, 90b, 90a ', 90b ' and a single Hπ when their angle of inclination α is determined automatically minimum value of the ratio of L _G and n _G is determined I understand that.

  Further, Expression 15- (31) is obtained from Expression 15- (19), (24), (25).

When n _G and α are determined from Equation 15- (31), n _C34 = the minimum value of the line segment C ₀ C ₃₄ is obtained. The form when the center of gravity G ₀ is at the midpoint of the line segment C ₀ C ₃₄ and when the first expression of Expression 19- (5) is substituted into the second expression and Expression 15- (31) has the form If you think that they are exactly the same, you can say:

  By substituting the first equation of equation 19- (5) into the second equation, the relationship of equation 15- (32) is obtained.

  Further, the relationship of Equation 15- (33) is obtained from Equation 15- (31).

  If α = 0 in Expression 15- (33), Expression 15- (33) is the same as Expression 15- (32).

Even if the propulsive force of one cylindrical propulsion unit among the four cylindrical propulsion units 90a, 90b, 90a ′, 90b ′ is changed by Δf _P from the equations 15- (32) (33), It can be seen that this results in the problem of runout of the rotating shaft when one propeller is regarded as one propeller (hereinafter referred to as a virtual coupled propeller).

Therefore, even if the propulsive force of each cylindrical propulsion unit 90a, 90b, 90a ′, 90b ′ of the cross-shaped connecting unit 110B is slightly increased or decreased, the swing angle of the rotating shaft of the general virtual propeller only changes slightly. In order to maintain the stability of the airframe 110B against such increase and decrease of the propulsive force, the ratio of L / n _G should be set sufficiently large, that is, n _C34 (in other words, the two rear parts of the cross-shaped connecting unit 110B have It can be seen that the distance between the flow wind pressure center points) should be set sufficiently long.

  A radial 8-connection unit 110C in which the above-described eight cylindrical propulsion units are connected in a radial manner, or a radial n-connection unit in which a larger number of cylindrical propulsion units are connected in a radial manner, represented by Formula 15- (30) or Formula 15- If the aircraft is designed to fully satisfy (33), the aircraft will be able to show hovering that is completely stable, does not move against the outside wind, and offsets the propeller anti-torque.

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).

  For example, if the 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 is significantly 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.

  If the number of the radial stabilizing blades 93 and 94 and the number of the cylindrical stabilizing 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.

  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.

  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 18 using the present invention, the realization of a flying car is no longer a dream.

S19. (Conditions to stabilize the aircraft by offsetting the effects of the propeller rotation shaft run-out)
FIG. 15 is a view of a certain moment when the propeller rotates, for example, in the airframe 80 of FIG. 12 when viewed from the side. In FIG. 15, point O is a fixed point of the propeller rotation axis, point P is an intersection of a horizontal line including the propeller operating point (center point) Y and a vertical line including point O, and point G ₁ is , The center of gravity when the center of gravity G of the body 80 is below the point O, and the point G ₂ is the position of the center of gravity when the center of gravity G of the body 80 is between the points O and P. G ₃ is the position of the center of gravity when the center of gravity G of the body 80 is above the point P. Further, the n value of the distance between the point O and the center of gravity G of the body 80 is n _G.

Here, a moment balance equation around a horizontal line passing through the center of gravity G of the body 80 when the center of gravity G of the body 80 is at each point G _{1 to} G ₃ will be considered.

(1) the moment of this case when the center of gravity G of the machine body 80 is at point G _1, towards the moment of the horizontal component of the thrust F _P of the propeller 81 (F _p sin .beta) is the vertical component of the F _P (mg) Since the moments are opposite to each other, the moment balance equation is expressed by Equation 19- (6).

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.

Since F _P = mg / cos β, F _C = Hπ mg tan β, and F _W = _W mg tan β, substituting these equations into Equation 19- (6) yields Equation 19- (1).

(2) When the center of gravity G of the airframe 80 is at the point G _{2 In} this case, the magnitudes of the moments of the horizontal component (F _p sin β) and the vertical component (mg) of the propeller propulsive force F _P are the above (1 However, considering the moment balance equation, Equation 19- (1) is obtained in the same manner as (1) above.

(3) When the center of gravity G of the airframe 80 is at the point G _{3 In} this case, the directions of the moments of the horizontal component (F _p sinβ) and the vertical component (mg) of the propeller thrust force F _P are the same. As in the case of the above (1), Expression 19- (1) is obtained.

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.

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.

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).

  Therefore, the condition for the airframe 80 to stably hover can be expressed as in Expression 19- (5).

  When the airframe 80 is stably hovering, it can be said that Expression 19- (5) is always satisfied.

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. (Fixed position of wind pressure center point regardless of propeller position in cylindrical stabilizer blade)
For example, in the fuselage 80 of FIG. 12, the length of the radial stabilizer wing 82 in the cylindrical stabilizer wing 83 is shortened to, for example, about half, and the position of the propeller 81 together with the radial stabilizer wing 82 is set in the vertical direction of the cylindrical stabilizer wing 83. An experiment was performed in which the center of gravity position at which the airframe 80 stably hovered was obtained at each position.

  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.

  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.

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 where 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 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.

  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.

  In this case, the auxiliary stabilizer wing 83b may be attached to the upper side of the airframe 80 instead of below it. However, when attached to the upper side, the wind pressure applied to the auxiliary stabilizer wing 83b is applied to the auxiliary stabilizer wing 83b when attached below. 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.

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.

  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.

S21. (Horizontal flight with a cylindrical propulsion unit)
FIG. 20 is a diagram showing a state in which an aircraft (for example, 90) composed of one cylindrical propulsion unit is flying horizontally at an inclination angle (an angle with respect to the vertical direction of the central axis of the aircraft) α (β with respect to the horizontal line). It is. And level flight speed and v, the propeller slipstream velocity and v _P, a propeller propulsion and F _P, the propeller slipstream wind (lift) and F _C, the external wind pressure and F _W.

Here, the terminal speed when the horizontal flight speed is gradually increased while the angle of attack of the airframe 90 in FIG. 20 is fixed will be considered. Now, an acceleration in the horizontal direction component of the propeller thrust F _P and a _1, the mass of the aircraft 90 and m _1, as further advances the angle of attack of the aircraft 90 in a horizontal direction while fixing the beta, the body 90, Consider a car fitted with a vehicle body 200 having wheels 200a. However, the wheel 200a of the vehicle body 200 has no friction at all, and the mass of the vehicle body 200 is 0 (zero) when running on the floor, and when it rises from the floor, the combined weight of the m ₂ (this situation, it is necessary when considering the thrust amplification device, such as in FIG. 10.).

It is assumed that no air force such as air resistance acts on the body 200 other than the airframe. All buoyancy this total body occurs stable wing aircraft 90 by proceeding in a horizontal direction and F, to the acceleration and a _2. It is assumed that the total body 210 of the body 90 and the body 200 is moving in the horizontal direction at the terminal speed v. In the above situation, Expression 21- (1) is established.

Here, assuming that the force applied to the stable blade by changing the direction of the propeller wake velocity v _P by the velocity v is Δf, Equation 21- (2) is established from FIG.

  Further, Expressions 21- (3) to (5) are established.

  The buoyancy F can also be expressed as shown in Equation 21- (6).

Substituting Equations 21- (1) to (5) into Equation 21- (6) and setting π (1-sinβ _W ) = Δπ, Equation 21- (8) is obtained.

  Further, the horizontal driving force f is expressed by Equation 21- (9).

  By substituting the formulas 21- (1) to (4) into the formula 21- (9), the formula 21- (10) is obtained.

When f = 0, since the united body 210 will proceed at a terminal velocity v, Motomari the relationship of Equation 21 (14), the ratio v / v _P of the terminal velocity v and the propeller slipstream velocity v _P , Β and the relationship between (HΔπ + W).

  Substituting equations 21- (12) to (14) into equation 21- (8) yields equation 21- (15), and transforming equation 21- (15) yields equation 21- (16).

And since the denominator (a ₁ / cosβ) on the left side of Equation 21- (16) corresponds to a ₀ when placed as Equation 21- (17), Equation 21- (16) becomes Equation 21- ( 18) It can be expressed as (19).

From equation 21 (19), if the aircraft 90 is in level flight at angles of attack beta, value sum (lift) F buoyancy when it reaches the terminal velocity v, obtained by dividing the original propeller thrust F _P in sinβ I understand that It can be seen that the magnitude of the lift F is related only to the angle of attack β. Here, in order to float the united airframe 210, a ₂ ≧ g must be satisfied. Therefore, the relationship of Equation 21- (18) to Equation 21- (20) is obtained. propeller propulsion force F _P that can be carried out can be seen.

The relationship of Equation 21- (22) is obtained from Equation 21- (2) (11). From this relation, if the 20 of the aircraft 90 is equal Hayami flat flight at angles of attack beta, the sum of the forces exerted on each stable wing, it is understood that a value obtained by dividing the propeller thrust F _P in tan.

If the horizontal component of F _P is F _h , Expression 21- (23) is obtained.

Further, since the lift L of a normal airplane is L = F _C + F _W and the propellers of a normal airplane are oriented horizontally, the lift L is given by Expression 21- (24) from Expressions 21- (22) and (23). ) From this equation, lift L becomes a horizontal thrust F _h divided by the sin .beta.

From Expression 21- (12), it can be seen that the higher the value of HΔπ + W, the smaller the terminal flight velocity v of the horizontal flight of the aircraft 90. Therefore, by setting the value of the F _P Formula 21 than the value represented by (20), if by increasing the value of HΔπ + W as possible, the sliding distance to the aircraft 90 is flying, attack It can be made very short by adjusting the angle β.

  Here, the two cylindrical propulsion units 90a and 90b of the fuselage 130 in FIG. 21 are coupled to the shaking of the fuselage when they are connected to each other in the direction of their wind flow (the direction of their central axes) as in the fuselage 130B in FIG. Consider stability. The airframe 130B of FIG. 23 is obtained by connecting two cylindrical propulsion units 90a and 90b while being shifted from each other in the wind flow direction in parallel with each other (this connection is referred to as oblique connection).

Note that the points P _i , C _i , G _i , W _i , H _i , and m _i (i = 1, 2) in FIG. 23 are propulsion force action points (propeller force action points) of the cylindrical propulsion units 90a and 90b, respectively. ), Wake wind pressure center point, center of gravity, external wind pressure center point, H value, and mass. Points P ₀ , C ₀ , G ₀ , and W ₀ are an overall driving force action point, an overall wake wind pressure center point, an overall center of gravity, and an overall external wind pressure center point of the airframe 130B, respectively. Point C ₁ ′ is the intersection of the perpendicular from the point C ₁ to the central axis 130a and the central axis 130a, and point C ₂ ′ is the intersection of the perpendicular from the point C ₂ to the central axis 130a and the central axis 130a. It is. Central axis 130a, the overall center of gravity G ₀ street respective tubular propulsion unit 90a, a straight line parallel to 90b the central axis of the (cylindrical shaft). Each point P _i , C _i , G _i , W _i (i = 1, 2) is arranged on the central axis of the cylindrical propulsion units 90a, 90b, respectively.

In FIG. 23, each cylindrical propulsion unit 90a, 90b is connected to each other by a rod-like connecting member 130b. The connecting member 130b is arranged along a straight line passing through the gravity centers G ₁ and G ₂ of the cylindrical propulsion units 90a and 90b.

It is assumed that each of the cylindrical propulsion units 90a and 90b has propulsive forces F _P1 and F _P2 that support their own weights m ₁ g and m ₂ g. That is, it is assumed that Expression 21- (25) is established.

Therefore, when the forces of F _C applied to the downstream wind pressure central points C ₁ and C ₂ of the cylindrical propulsion units 90a and 90b are F _C1 and F _C2 , Expressions 21- (26) and (27) are established.

Since each cylindrical propulsion unit 90a, 90b has propulsion forces F _P1 , F _P2 that support its own weights m ₁ g, m ₂ g, the total center of gravity G ₀ of the fuselage 130B is determined by each cylindrical propulsion unit 90a, The length of the line segment G ₁ G ₂ is proportionally divided by the ratio of m ₁ and m ₂ on the straight line connecting the centroids G ₁ and G ₂ of 90b. Further, the total thrust force action point P ₀ of the thrust force action points P ₁ and P ₂ is on the total gravity center G ₀ .

In FIG. 23, the points P ₀ and G ₀ are located on the central axis 130a, but since H ₁ ≠ H ₂ , the total wake wind pressure central point C ₀ is not located on the central axis 130a. Obviously, when the points P ₀ , C ₀ , G ₀ are not on the central axis 130a, the airframe 130B becomes unstable.

  Here, Expression 21- (28) is established from Expression 21- (25), and Expression 21- (29) is established from Expression 21- (26) (27).

Therefore, from Expressions 21- (28) and (29), the condition for arranging the points P ₀ , C ₀ , G ₀ on the central axis 130a is Expression 21- (30).

  Therefore, from Equation 21- (30), it can be seen that one of the conditions for the airframe 130B to be stable is Equation 21- (31).

Similarly, F _W1 and F _W2 are the forces F _W applied to the points W ₁ and W ₂ of the cylindrical propulsion units 90a and 90b, and W (= S _W / S _C ) of the cylindrical propulsion units 90a and 90b. When the values are W ₁ and W ₂ , Expressions 21- (32) and (33) are established.

Therefore, from the expressions 21- (32) and (33), the condition for arranging the points P ₀ , C ₀ , G ₀ on the central axis 130a is expressed by the expression 21- (34).

Then, the total driving force applied to the vehicle body 130B and F _P, the F _C becomes the force and F _C, the overall external wind pressure and F _W, each tubular propulsion unit 90a, the relationship between the H values and W values of 90b H ₁ = H ₂ = is H and W ₁ = W ₂ = W, when put F _P as of formula 21 (35), wherein the formula 21- (26) (27) (32) (33) 21 (36) Obtain (37).

Then, when n _G1W1 / n _G1C1 , n _G2W2 / n _G2C2 , n _G0W0 / n _G0C0 are obtained using Expressions 21- (35), (36), and (37), Expression 21- (38) is obtained and each point P _{0 is obtained.} , C ₀ , G ₀ are located on the central axis 130a.

From Equation 21- (38), if the relationship of H ₁ = H ₂ = H and W ₁ = W ₂ = W holds for the H value and W value of each cylindrical propulsion unit 90a, 90b, each cylindrical propulsion unit Regardless of the masses m ₁ and m ₂ of 90a and 90b, the points P ₀ , C ₀ and G ₀ are located on the central axis 130a, and the airframe 130B is stabilized.

Furthermore, a resistance moment is generated when the airframe 130B rotates due to the counter-torque of the propeller rotation by the forces F _C1 and F _C2 generated at the points C ₁ and C ₂ of the two cylindrical propulsion units 90a and 90b that are independent from each other. Therefore, coupled with the counter-torque canceling force of the in-cylinder radial stabilizer blades of the cylindrical propulsion units 90a and 90b, the forward and reverse rotation of the airframe 130B is almost eliminated.

In addition, F _C1 and F _C2 applied to the points C1 and C2 of the two cylindrical propulsion units 90a and 90b that are independent from each other are actually each point C with respect to the back-and-forth and right-and-left swings around the center of gravity G _{0. Since} it is the same as that applied to ₁ ′ and C ₂ ′, the resistance moment to the above-described back-and-forth and left-right swing becomes very large.

  Therefore, the stability of the fuselage 130B is greatly improved compared to a single cylindrical propulsion unit.

Further, from the relationship of W value W ₁ = W ₂ , it is clear that the cylindrical diameters of the cylindrical propulsion units 90a and 90b must be the same, so the sizes of the cylindrical propulsion units 90a and 90b and the respective condition values (each It can be seen that the points C ₁ and C ₂ and the points G ₁ and G ₂ (except for the positional relationship and masses m ₁ and m ₂ ) must be exactly the same.

Further, the smaller the angle α formed between the connecting member 130b of the machine body 130B and the central shaft 130a, the longer the line segment G ₀ C ₁ ′ and the line segment G ₀ C ₂ ′, and the resistance moment due to F _C1 and F _C2 increases. It can also be seen that the stability of the airframe 130B increases.

Here, even if an auxiliary wing and a weight are added to the fuselage 130B and the point W ₀ and the point G _{0 of} the fuselage 130B are adjusted so as to coincide with the point C ₀ , if Formula 21- (39) is satisfied. The airframe 130B is stabilized.

If the distance between the points G ₀ and C ₁ ′ is n _{G0C1 ′} and the distance between the points G ₀ and C ₂ ′ is n _{G0C2 ′} , then n _{G0C1 ′} = n _{G0C2 ′} , so that n _{G0C1 ′} = If n _{G0C2 ′} = n _GC and n _G0P0 = n _G , then Equation 21- (39) becomes Equation 21- (40). This equation is the stability condition itself against the shaking of the airframe 130B.

  Next, as shown in FIG. 24, the stability with respect to shaking in the airframe 150 in which three cylindrical propulsion units (for example, 90) are obliquely connected will be examined.

Note that the points C _i , G _i , W _i , H _i , and m _i (i = 1, 2, 3) in FIG. 24 are respectively the wake wind pressure center point, the center of gravity, and the center of gravity of the cylindrical propulsion units 90a, 90b, 90c. The external wind pressure center point, H value, and mass. Here, H ₁ = H ₂ = H and W ₁ = W ₂ = W.

In the airframe 150, the cylindrical propulsion units 90a and 90b are obliquely connected in parallel to each other by, for example, a rod-shaped connecting member 150a. The connecting member 150a is arranged along a straight line passing through the gravity centers G ₁ and G ₂ of the cylindrical propulsion units 90a and 90b. Further, the cylindrical propulsion unit 90c is connected to the connection member 150a by, for example, a rod-shaped connection member 150b, so that the cylindrical propulsion units 90a and 90b are connected so as to be parallel to them at the rear, for example. The connecting member 150b is an intermediate point of the connecting member 150a (more specifically, a point obtained by dividing the length of the line segment G ₁ G _{2 by} the ratio of m ₁ and m ₂ on the straight line connecting the centroids G ₁ and G _2. ) are arranged along a straight line passing through the center of gravity G3 of the G ₁₂ and the tubular propulsion unit 90c.

In addition, 150c in the figure is a virtual connection unit in which the oblique connection unit of each of the cylindrical propulsion units 90a and 90b is regarded as a single unit. Point G ₁₂₃ is a point obtained by dividing the length of line segment G ₁₂ G _{3 by} the ratio of (m ₁ + m ₂ ), m ₃ on the straight line connecting points G ₁₂ and G ₃ . Point C ₁₂₃ is a point on the side of point C ₁₂ when line segment C ₁₂ C ₃ is divided into three equal parts.

First, assuming that the two oblique connection units of the cylindrical propulsion units 90a and 90b are stably hovering, when considering the virtual connection unit 150c in which the two oblique connection units are regarded as a single unit, the virtual connection unit 150c. propulsion F _P ', F _C becomes the force F _C' and an external wind pressure F _W 'is a formula 21- (35) (36) (37) from equation 21 (41) (42) (43) .

  When another tubular propulsion unit 90c is obliquely connected to the two oblique connection units, the condition for stable hovering of the whole (that is, the airframe 150) is a tubular shape added to the virtual connection unit 150c. This results in the problem of the two diagonal connection of the two propulsion units with the propulsion unit 90c. Therefore, it can be seen that the stability of the airframe 150 cannot be ensured unless the H value, W value, etc. of the third added cylindrical propulsion unit 90c are the same H and W as the cylindrical propulsion units 90a, 90b.

From the above, in order to connect and stabilize a plurality of cylindrical propulsion units, the stability condition values of the respective cylindrical propulsion units other than the mass relationship of all the cylindrical propulsion units and the positional relationship between the center point of the wake wind pressure and the center of gravity ( It can be seen that all of the values (H value, W value, r ₀ value, etc.) must be made equal (this is called the 2-diagonal connection theory). And when considering the stability of these cylindrical propulsion units, they can be considered as a stability problem when assuming a general virtual unit at the position of the total wake wind pressure center point of each wake wind pressure center point. Yes (however, the propulsive force and mass are the total value of the plurality of cylindrical propulsion units).

An airframe 130B shown in FIG. 23 was actually made, and an experiment was conducted to confirm whether the above-mentioned two-diagonal connection theory is correct. The H value, W value, and r _{0 of} these cylindrical propulsion units 90a and 90b (referred to as the attribute condition values of the cylindrical propulsion unit) are H = 7.388, W = 0.8546, r ₀ = 15, respectively. The airframe 130B is constructed by setting the distance between the propeller rotating shafts of each cylindrical propulsion unit to 37 cm, and further setting the vertical displacement length of each cylindrical propulsion unit to 20 cm (however, each cylinder When the n value of the length of the in-cylinder cross stabilizer in the cylindrical propulsion unit is set to 1.44 so that the counter-torque caused by the propeller rotation of each cylindrical propulsion unit can be almost offset), the airframe 130B is hovered. Despite the propellers rotating in the same direction, the airframe 130B did not rotate counterclockwise or rotate in the forward direction, and performed stable hovering with no left-right swing.

  From the above experiment, when the separate cylindrical propulsion units 90a and 90b having independent propulsion wind flows are obliquely connected as shown in FIG. 23, if the attribute condition values of these cylindrical propulsion units are the same, the airframe 130B It was confirmed that the rotation tendency due to the rotation of the entire propeller disappeared, and the stability against left and right runout increased, and the above-mentioned two-angled connection theory was proved to be correct. It is obvious that this can be applied not only to the two diagonal connection but also to the case where two or more cylindrical propulsion units are similarly diagonally connected.

  23 and 24, the conditions for connecting and stabilizing a plurality of cylindrical propulsion units are as follows: the overall thrust force action point of the entire fuselage, the overall center of gravity, the overall wake wind pressure center point, and the overall external wind pressure center point Are all aligned on the same straight line (central axis of the entire aircraft) parallel to the central axis of each cylindrical propulsion unit.

  Therefore, in the airframe 140 of FIG. 22, regarding the propulsion unit 140c, as long as the other two propulsion units 140a and 140b have the same attribute condition value, the attributes different from those of the other two propulsion units 140a and 140b. It can be seen that the entire airframe 140 stably hovers even if it has the condition value.

The total center of gravity G ₀ when performing the most stable hovering in the experiment of the above-described two-slanted connecting unit 130B was when it coincided with the total wake wind pressure center point C ₀ . What does this mean? In the case of the two oblique connecting unit 130B, considering that the total center of gravity G ₀ is located at the midpoint between the wake air pressure center points C ₁ and C _{2 of} the two independent cylindrical propulsion units 90a and 90b, the equation 21 − It can be seen that (40) holds. However, although the total external wind pressure center point W ₀ did not coincide with the total center of gravity G ₀ , the fuselage 130B was stably hovering. This means that stable hovering can be performed even if the first expression of Expression 19- (5) is violated.

Each wake wind pressure center point of the n _GC airframe 130B C of formula _{21- (40) 1 ', C} 2' the distance between the overall center of gravity G ₀ and n _{G0C1 ',} n _G0C2' formula rewriting was replaced by the expression 21 -Considering (39), the following can be understood.

The airframe 160 in FIG. 25 is configured by obliquely connecting four cylindrical propulsion units (for example, 90) having the same attribute condition value. FIG. 25 shows the wake wind pressure center point C _i and the external wind pressure center point W _i (i = 1 to 4) of each cylindrical propulsion unit 90a to 90d, the overall wake wind pressure center point C _{0 of the} fuselage 160, and the overall exterior. it is a diagram showing the positional relationship between the wind pressure center point W _0.

Point C ₁₂ is a midpoint of line segment C ₁ C ₂ , and point C ₁₂₃ is a point on the side of point C ₁₂ when line segment C ₁₂ C ₃ is divided into three equal parts. C ₀ is the point closest to the point C ₁₂₃ when the line segment C ₁₂₃ C ₄ is divided into four equal parts. The total center of gravity G ₀ is adjusted to coincide with the point C ₀ . The total external wind pressure center point W ₀ is also determined in the same manner as the point C ₀ . Then, considering a straight line VV ′ obtained by extending the line segment G ₀ W ₀ , the intersection point between the perpendicular line from the wake wind pressure center point C _i to the straight line VV ′ and the straight line VV ′ to each cylindrical propulsion unit 90a to 90d is represented by C _i ′. And The attribute condition value of each tubular propulsion units 90a to 90d H, is W, the n value of the distance between the overall propeller shaft fixed point and the point G ₀ of the virtual connecting unit of each tubular propulsion units 90a to 90d n _G And the n value of the distance between each point C _i ′ (i = 1 to 4) and the point G ₀ is n _{GCi ′} and the n value of the distance between the point G ₀ and the point W ₀ is n _GW . Further, the propulsive force of each of the cylindrical propulsion units 90a to 90d is set to F _P1 = F _P2 = F _P3 = F _P4 . The hovering stability condition of the airframe 160 is considered to be Formula 21- (44) -1. Note that “4” on the right side of Expression 21- (44) -1 is the number of cylindrical propulsion units (90a to 90d in this example) having the same attribute.

W in Expression 21- (44) -1 is exactly the same as the W value of the entire aircraft including the connecting member. Here, Expression 21- (44) -1 when F _P1 ≠ F _P2 ≠ F _P3 ≠ F _P4 becomes Expression 21- (44) -2.

Furthermore, Formula 21- (44) -2 in the case of the fuselage 161 in which the fuselage 90 ′ of FIG. 25 is added to the fuselage 160 along the central axis VV ′ of the fuselage 160 becomes Formula 21- (44) -2 ′. . However, the total center of gravity G ₀ is slightly lower than the position of FIG.

The general formula when the number of propulsion units other than the propulsion unit 90 ′ of the airframe 161 is increased is expressed by Formula 21- (44) -3. In FIG. 25, symbol F _P ′ is the propulsive force of the fuselage 90 ′, symbol W ′ is the external wind pressure center point of the fuselage 90 ′, symbol C ′ is the wake wind pressure center point of the fuselage 90 ′, H ′ is the attribute value H of the aircraft 90 ′.

  Expression 21- (44) -3 can be said to be a basic general conditional expression for performing stable hovering when a plurality of propulsion units are obliquely connected.

  When considering the stability of the airframe with a single cylindrical propulsion unit, the position of the downstream wind pressure center point may be adjusted to stop the airframe shake due to the propeller rotation shaft runout, but the downstream wind pressure center point Is displaced from the center of gravity, and another moment force applied to the external wind pressure center point that balances the moment force of the subsequent wind pressure center point is required. The conditional expression was Expression 19- (5).

  However, in the case of a fuselage in which a plurality of cylindrical propulsion units (not limited to a cylindrical shape) are obliquely connected as shown in FIGS. 22 to 25, there are a plurality of independent wake wind pressure center points. Place the center of gravity at the total wake wind pressure center point of the multiple wake wind pressure center points so that the wake wind pressure applied to multiple wake wind pressure center points balances the center of gravity of the aircraft, and the moment by each wake wind pressure It can be seen that the aircraft performs stable hovering if the sum of the values is sufficiently larger than the sum of the moments caused by the propeller rotation shafts of the virtual connecting units of all the cylindrical propulsion units and the moment caused by the total external wind pressure. At that time, in the case of having a plurality of cylindrical propulsion units, the larger the value on the left side of Equation 21- (44) -3, the more the influence of the external wind pressure is eliminated in balance, and the aircraft can stably hover. It turns out that it becomes easy to do.

  The following can be considered as problems when the aircraft having a single cylindrical propulsion unit flies horizontally. When considering the situation where the aircraft starts level flight and gradually increases in speed, a constant speed propulsion wind flows on a stable wing within the range of the propulsion wind flow flowing from the cylindrical propulsion unit. On the stable wing outside the range, the external wind flow from the front of the fuselage flows while gradually increasing its speed, and the wind pressure of the external wind flow gradually increases. Due to this increase, the balance of the total wind flow pressure applied to the aircraft (the sum of the wake wind pressure and the external wind pressure) is lost, and the aircraft becomes unstable.

  In order to solve this problem, it is necessary to provide a plurality of cylindrical propulsion units (generally wind flow generating devices) of the airframe. Then, the overall external wind pressure center point of the stable wing outside the range of the propulsion wind flow generated from the cylindrical propulsion unit (this is called the self wind flow) is changed to the total wake wind pressure center point of the stable wing within the self wind flow range. It is to match, and to match the total center of gravity to that point. Of course, it is an indispensable condition that a plurality of cylindrical propulsion units are obliquely connected as described above. In addition, if each cylindrical propulsion unit is arranged so as to satisfy Equation 21- (44) -3, it is natural that the aircraft will stably hover, but even when starting horizontal flight. You can see that the flight is very stable.

  In addition, the higher the speed, the stronger the external wind pressure applied to the stable wings outside the range of the self-wind flow of each cylindrical propulsion unit. As a result, the stability of the entire aircraft increases more and more.

  Furthermore, since such an airframe is not affected by the external wind pressure, stable horizontal flight can be performed without a normal wing. Furthermore, such a fuselage can be used in a vacuum, more specifically in space, by replacing a propeller-type cylindrical propulsion unit (wind flow generator) with an injection-type propulsion unit (for example, a gas injector or a rocket injector). In other words, it can be seen that stable hovering and horizontal flight can be performed even in the attractive range of other celestial bodies without air.

  Equation 19- (5) is a stable hovering condition for a fuselage having a single cylindrical propulsion unit, and a stable hovering and stable horizontal flight condition formula for a fuselage having a plurality of independent cylindrical propulsion units is: Equation 21- (44) -3 is obtained.

  If Expression 21- (44) -3 is satisfied, it can be seen that the airframe 110C of the radial 8-connected unit in FIG. 18 can perform stable hovering even if the external wind pressure center point does not coincide with the center of gravity.

  Next, as a specific method for removing the instability due to the external wind flow from the front during horizontal flight in the case of an airframe in which a plurality of cylindrical propulsion units are obliquely connected, consider, for example, the airframe 160 of FIG.

  FIG. 26 is a view of the airframe 160 seen from above, which is flying horizontally upward on the paper surface.

  Airframe 160 is configured by connecting three cylindrical propulsion units (for example, 90 (90a, 90b, 90c)) having the same attribute condition value. The cylindrical propulsion units 90a and 90b are connected to each other in parallel with a space therebetween by, for example, a rod-like connecting member 160a. For example, a rod-like connecting member 160c is obliquely connected to each of the cylindrical propulsion units 90a and 90b so as to be disposed on 130b.

  The airframe 160 includes, for example, a cylindrical (or radial) auxiliary wing 160e. The auxiliary wing 160e is obliquely connected to each of the propulsion units 90a and 90b by, for example, a rod-shaped connecting member 160d so as to be disposed coaxially with the central shaft 160b of the entire cylindrical propulsion unit 90a and 90b, for example. ing. The diameter (or width and height) of the auxiliary wing 160e is set to the same dimension as the diameter of each cylindrical propulsion unit 90a, 90b, 90c.

  In addition, the code | symbol 160f of FIG. 26 has shown the virtual connection unit at the time of considering the three cylindrical propulsion units 90a, 90b, 90c as a single unit.

For each of the cylindrical propulsion units 90a, 90b, 90c, the wake wind pressure center points thereof are C ₁ , C ₂ , C _3, and the external wind wind pressure center points of the wind flow flowing outside the outer tubes are C _{h 1} , C _h2 and _Ch3, and the outer H values of the outer cylinders are set equal to H. Regarding the auxiliary wing 160e, the center point of the external wind flow pressure due to the wind flow from the front is C _X and the H value is H _X. Further, regarding the virtual connection unit 160f, the total wake wind pressure center point is C ₀ and the H value outside the outer cylinder is H ₀ .

In the airframe 160, the total center of gravity G ₀ is made to coincide with the point C ₀ .

Here, the positions and lengths of the radial stabilizer blades and the in-cylinder cylindrical stabilizer blades in the respective cylindrical propulsion units 90a, 90b, 90c are adjusted, and the wind flow wind pressure center points C _i , If C _hi (i = 1 to 3) are matched with each other, the wind current and wind pressure central points C ₀ and C _{h0 of} the virtual connection unit 160f are also matched with each other. Since the overall center of gravity G ₀ coincides with the point C ₀ , the overall external wind pressure center point C _h0 coincides with the overall center of gravity G ₀ , so that no matter what speed the aircraft 160 is flying at, The instability of 160 no longer occurs. Under such a setting, the auxiliary wing 160e is not necessary. However, in reality, it is expected that the wind flow and wind pressure center points C _i and C _hi of the cylindrical propulsion units 90a, 90b, and 90c often do not coincide with each other. Further, it is expected that the airflow and wind pressure center points C ₀ and C _h0 of the airframe 160 often do not coincide with each other depending on the size and shape of the connecting members 160a, 160c, and 160d. FIG. 26 assumes that the points C ₀ and C _h0 do not coincide with each other, so that the instability of the airframe 160 is not generated by providing the auxiliary wing 160e.

Here, if the distance between the general external wind pressure central point C _h0 and the total gravity center G ₀ is n _Gh , and the distance between the external wind wind pressure central point C _{X of the} auxiliary wing 160e and the total gravity center G ₀ is n _GX , In order to make the total wind pressure center point of C _X and the point C _h0 coincide with the point G ₀ , it is obvious that Expression 21- (45) should be satisfied.

H ₀ can be regarded as the total value of the H values of the cylindrical propulsion units 90a, 90b, and 90c, and therefore can be expressed as Expression 21- (46).

  Therefore, from Expression 21- (46), Expression 21- (45) becomes Expression 21- (47).

  Therefore, if the auxiliary wing 160e is attached so as to satisfy Expression 21- (47), a moment that cancels out the wind pressure moment due to the wind flow from the front is generated, so that the entire body 160 can perform a stable horizontal flight.

  Next, when the airframe 160 (airframe in which a plurality of cylindrical propulsion units are obliquely connected) 160 is flying in the horizontal direction, in order to cancel the moment that the airframe 160 tries to return to the vertical posture by the restoring effect (ie, In order for the airframe 160 to make a stable transition from hovering to level flight), what should be done?

  Before that, consider an airframe (for example, airframe 180 in FIG. 27) composed of, for example, a cylindrical propulsion unit with a single wind flow.

  If the airframe 180 of FIG. 27 is stably flying in the horizontal direction, the moment balance equation at this time is usually Equation 21- (48). For the sake of simplicity, the external wind flow that flows outside the outer cylinder is ignored (because the final conclusion is the same even if this external wind flow is considered).

However, since the moment on the left side of Equation 21- (48) is a couple moment, Equation 21- (48) is correctly a couple moment equation applied to the resultant action point of F _C and F _W. (49).

  Here, Equations 21- (50) and 51 are obtained by Equations 21- (2), (3), and (4).

  Further, when Expressions 22- (50) and (51) are substituted into Expression 21- (49), Expression 21- (52) is obtained.

  Since Expression 21- (12) becomes Expression 21- (12), Expression 21- (53) is obtained.

  The meaning of Equation 21- (53) is that the point C is in front of (above) the point G, and the first equation of Equation 19- (5) must be satisfied.

  Also, from equation 21- (53), it can be seen that even in equation 21- (54), the airframe 180 of FIG. 27 can perform horizontal flight while supporting the gravity after offsetting the restoring effect of the gravity.

  Therefore, in FIG. 27, it is necessary to avoid that the total wind current center point C moves behind (below) the center of gravity G in order to cancel the restoration effect. When the point C moves after (below) the point G as shown in FIG. 27, a reverse moment more than the moment of the gravity restoring effect is applied too much and the body 180 turns downward.

  As shown in FIG. 20, when the total wind current center point C is in front of (above) the total center of gravity G, the equation corresponding to the equation 21- (52) for canceling the effect of restoring gravity is similarly the moment balance equation. When calculated more, Equation 21- (55) is obtained.

  Expression 21- (53) becomes Expression 21- (56).

  The difference between Equation 21- (56) and Equation 19- (5) is the difference between Δπ and π on the left side. This is particularly important in order to make a stable transition when the fuselage (for example, the cylindrical propulsion unit) 180 is shifted from the hovering state to the horizontal flight state. Obviously, in order to eliminate the influence of this difference, it is sufficient to satisfy Expression 21- (54). That is, the overall wake wind pressure center point C, the overall center of gravity G, and the external wind pressure center point W of the airframe 180 are matched. By matching the points C, G, and W, the airframe 180 can perform stable hovering and smoothly transition to horizontal flight (however, it is necessary to have n oblique connections by a plurality of wind currents).

  Substituting Equation 21- (54) into Equation 21- (55) yields Equation 21- (57). This equation is the same as Equation 21- (12).

  Therefore, from Expression 21- (57), the horizontal movement speed v becomes Expression 21- (58). If the angle β is gradually decreased from 90 °, the horizontal movement speed v of the airframe 180 increases gradually from 0 accordingly. It will be a thing.

  When the airframe is composed of a plurality of wind propulsion units (ie, in the case of the airframe 160 in FIG. 26), the airframe 160 can perform stable hovering if Equation 21- (44) -3 is satisfied. I know that. In this case, since the influence of the external wind pressure is almost eliminated, it can be said that the condition under which the airframe 160 composed of a plurality of wind flow propulsion units can stably shift from the hovering state to the horizontal flight is Expression 21- (59).

In the above calculation when the aircraft 180 in FIG. 27 is flying horizontally, the influence of the external wind flow flowing outside the outer cylinder is ignored. In this case, the influence of the external wind flow is taken into consideration. In this case, F _C in the expression 21- (50) becomes the expression 21- (60).

  When calculating using Expression 21- (60), Expression 21- (56) becomes Expression 21- (61).

  From Equation 21- (61), the condition for the aircraft 180 to maintain the horizontal flight by canceling the restoring effect due to gravity is to satisfy Equation 21- (54), as in the above conclusion.

  In the case of the fuselage 160 of FIG. 26, the total center of gravity, the total wake wind pressure center point, the total external wind pressure center point, and the total external wind flow wind pressure of the entire fuselage are also taken into consideration when considering the area and mass of the members connecting the propulsion units. If the center points coincide with each other and Expression 21- (44) -3 is satisfied, the aircraft can stably hover and move to level flight more stably.

S22. (Propulsion amplifier)
The airframe having a cylindrical propulsion unit floats the airframe by a propulsive force that is a fraction of the gravitational force applied to the airframe during horizontal flight and continues the horizontal flight as it is. Therefore, the airframe (propulsive force amplifying device) 170 in FIG. Considering that, the aircraft 170 can float with a propulsive force that is a fraction of the gravity of the aircraft 170.

The fuselage 170 has a plurality (for example, four) of n obliquely connected cylindrical propulsion units 170a to 170d, and these n obliquely connected cylindrical propulsion units are substantially symmetrical along a coaxial circumference (for example, the same circumference). And are connected to each other by a connecting member 170e so that their propulsive forces F _P1 , F _P2 , F _P3 , and F _P4 are directed in the same rotational direction.

  The connecting member 170e has, for example, a disk 170f, and is configured by forming a plurality of connecting portions 170g on the outer peripheral surface thereof, and n-tilt-connected cylindrical propulsion units 170a to 170d at the distal ends of the connecting portions 170g. Are connected.

For example, when the angle of attack β of each of the cylindrical propulsion units 170a to 170d in FIG. 10 is 5.73 °, the horizontal flight propulsion force F _Pi (i = 1 to 4) of each of the cylindrical propulsion units 170a to 170d is expressed by Equation 21. -From (19), Formula 22- (10) is obtained, and the airframe 170 is lifted with a propulsive force that is 1/40 times the gravity applied to the entire airframe 107.

  There is one problem with the propulsion amplifier 170 of FIG. When each cylindrical propulsion unit 170a to 170d is considered as two or more n oblique connection units, each cylindrical propulsion unit 170a to 170d is considered as one virtual unit representing the n oblique connection unit. At this time, there are two or more independent wind pressure center points of the virtual unit in the virtual unit, and by rotating around the center point of the disk 170f, each of the virtual units has its rotational moment. An anti-rotational moment with respect to As a result, the rotation of the entire body 170 is suppressed.

  Similarly, even when each of the cylindrical propulsion units 170a to 170d in FIG. 10 is considered as a single cylindrical propulsion unit, depending on the coupling position between the cylindrical propulsion units 170a to 170d and the disk 170, it may be the same as described above. A rotational moment is generated.

  FIG. 28 is a top view of the disk 170f of FIG. Each cylindrical propulsion unit 170a-170d is represented by a straight line. The cylindrical propulsion units 170a to 170d are arranged symmetrically so as to direct the propulsive force in the same rotational direction along the coaxial circumference even if their propeller rotational directions are the same. Anti-torque is offset.

  Each point P of FIG. 28 is a propeller action point of each cylindrical propulsion unit 170a-170b. Each of the cylindrical propulsion units 170a to 170d is connected to the disc 170f so as to coincide with the tangential direction of the disc 170f at its propeller action point P. Further, the positional relationship among the center of gravity G, the overall wake wind pressure center point C, and the external wind pressure center point W of each of the cylindrical propulsion units 170a to 170d is arranged so as to satisfy Expression 19- (5).

  In the propulsive force amplifying apparatus 170 configured in this way, the cylindrical propulsion units 170a to 170d are arranged substantially symmetrically so as to direct the propulsive force in the same rotational direction along the coaxial circumference. The instability of each of the cylindrical propulsion units 170a to 170d while rotating on the coaxial circumference is offset, and the safety of the entire body 170 is maintained.

Moreover, each cylindrical propulsion unit 170a-170d is connected to the disk 170f at their propeller action point P, and the positional relationship between these points G, C, W satisfies Expression 19- (5). alone stability of each tubular propulsion unit 170a~170d is maintained, anti-roll moment will occur commensurate with the rotational moment of the center of gravity around G proportional to the centripetal force mv ^2/2 received from the point P.

  In fact, when a cylindrical propulsion unit (for example, 90) used for the experiments so far is held around the propeller action point P and rotated around the experimenter's body, it can be seen that there is no resistance. If a similar experiment is performed with the hand around the center of gravity G other than the point P, for example, the counter-rotating force can be felt by the hand.

  It should be noted here that the external wind flow flowing outside the outer cylinder of each of the cylindrical propulsion units 170a to 170d gradually increases its speed, and as a result, is applied to the external wind flow pressure center point outside the outer cylinder. The wind pressure gradually increases.

  As a result, the stability of each of the cylindrical propulsion units 170a to 170d is lost, and as a result, a counter-rotation moment is generated again. In order to prevent this instability, the same method described with reference to FIG. 26 is used, and the moment balance around the center of gravity G of each cylindrical propulsion unit 170a to 170d flows outside the outer cylinder of each cylindrical propulsion unit 170a to 170d. It is clear that it should not be destroyed by wind currents.

When the propulsion amplifying device 170 of FIG. 28 or FIG. 10 starts to rise at the acceleration a ₂ represented by the equation 21- (19), the direction of the wind speed v hitting the stable blade in FIG. As a result, the ascending force itself decreases and the question arises that the ascending speed will slow down. However, when the direction of the speed v in FIG. 20 is changed downward, the propulsive force f of each of the cylindrical propulsion units 170a to 170d expressed by the equation 21- (10) becomes larger than 0, and f As the propulsion speed of each of the cylindrical propulsion units 170a to 170d increases until = 0, the ascending force of the propulsive force amplifying device 170 is maintained. This is proved by FIG. 29 and the equation as follows.

FIG. 29 is a diagram when the cylindrical propulsion unit 90 of FIG. 20 is rising upward at the speed v ₂ . Here, the horizontal speed is v ₁ . The ascending force F is expressed by Expression 22- (11) from Expression 21- (6).

Equation 22- (1) (2) (3) is established if the force of the wind speed v ₁ is Δf ₁ , the force of the wind speed v ₂ is Δf _2, and the wind pressure applied to the stable blade is Δf. To do.

  Here, Expression 22- (4) is obtained from Expression 21- (3) (4) (5) (6).

  Now, when Expression 22- (12) is put, Expression 22- (4) becomes Expression 22- (5).

  On the other hand, Expression 22- (6) is obtained from Expression 21- (9).

  Since the terminal speed is reached when f = 0, equation 22- (7) and equation 22- (8) are obtained by setting f = 0 in equation 22- (6).

When the formula 22-a (7) Substituting equation 22-a (5), the formula 22- (9), and this equation is exactly the same formula as formula 21 (19), rising force F F _P From this equation It turns out that it becomes the value which divided by sinβ.

Since the right side of Equation 21- (13) and the right side of Equation 22- (8) are the same, the wind pressure applied to the stable wing is proportional to v _P ² when the aircraft 90 is flying horizontally. , (HΔπ + W) and tan β, and if H, W, β, Δπ, and v _P are determined, it can be understood that the value is always constant regardless of the value of the ascending speed v ₂ of the aircraft.

S23. (Simultaneous rotation in the same direction around each center of gravity of each cylindrical propulsion unit with n diagonal connections)
FIG. 30 is a side view of an airframe 250 in which three, for example, cylindrical propulsion units (for example, 90) are connected obliquely, and the cylindrical propulsion units 90a to 90c are simplified and represented by straight lines. The attribute values of the three cylindrical propulsion units 90a to 90c are all equal. The reference numeral 250b in FIG. 30 is a connecting member that connects the cylindrical propulsion units 90a and 90b, and the reference numeral 250c is a connecting member that connects the cylindrical propulsion unit 90c and the connecting member 250b.

Each of the cylindrical propulsion units 90a and 90b is connected to the connecting member 250b so as to be rotatable around its center of gravity G ₁ and G ₂ . The tubular propulsion unit 90c, to the connecting member 250c, is pivotally connected around its center of gravity G _3.

Although being not shown, the body 250 has a respective tubular propulsion unit 90a~90c further comprising a rotating mechanism for simultaneously rotated in the same direction to each around the center of gravity G _1, G _2, G ₃ .

As described above, the three-diagonally connected machine body 250 can perform stable hovering. When each of the cylindrical propulsion units 90a to 90c is slowly rotated around the center of gravity G ₁ , G ₂ , G ₃ simultaneously in the same direction, for example, when each is rotated to an angle α A dotted line in FIG. 30 represents the state of each of the cylindrical propulsion units 90a to 90c. A central wavy line 250a represents a virtual unit representing each of the cylindrical propulsion units 90a to 90c.

As is clear from FIG. 30, even if each cylindrical propulsion unit 90a to 90c is rotated in the same direction by an angle α about the center of gravity G ₁ , G ₂ , G ₃ , the total center of gravity G of the three oblique connections is obtained. _It can be seen that the position of ₀ does not change. Virtual unit 250a of 3 oblique linked in the rotation of the tubular propulsion unit 90a~90c be understood that going to rotate around its center of gravity G ₀ in the same direction at the same time. Even after the rotation of the angle α, the cylindrical propulsion units 90a to 90c are connected in three diagonals, so that they remain stable with respect to rolling and counter-rotation.

  If each cylindrical propulsion unit 90a-90c inclines only the angle (alpha) like FIG. 30, naturally the whole body 250 will start a horizontal flight. In this case, if the auxiliary wing as described with reference to FIG. 26 is attached, the entire body 250 can perform a stable horizontal flight.

  However, in practice, in consideration of the area and mass of the connecting member when a plurality of propulsion units are connected, in order to perform stable transition to stable hovering and horizontal flight, as described above, Formula 21- (44) -3 when considering the entire aircraft must be satisfied. Furthermore, it is necessary to match the overall center of gravity of the entire aircraft, the overall wake wind pressure center point, the overall external wind pressure center point, and the overall external wind wind pressure center point. Further, in order to maintain the coincidence of these total points even when shifting to the horizontal transition, it is necessary to match the center of gravity of the downstream wind pressure and the center of gravity of each propulsion unit that is simultaneously rotated in the same direction. As long as the external wind pressure center point of each propulsion unit does not coincide with the wake wind pressure center point and the center of gravity, as long as Expression 21- (44) -3 is satisfied, the impact on the entire aircraft will be extremely small. It is.

As described above, if each of the cylindrical propulsion units 90a to 90c connected in an oblique direction is gently rotated in the same direction around the center of gravity G ₁ , G ₂ , G ₃ simultaneously, You can see that the transition to flight is very stable. If this is utilized, it seems that a flying car is fully feasible.

  FIG. 31 is an example of a flying car 300 that utilizes n-slanted connection and simultaneous rotation in the same direction. The flying car 300 includes a vehicle 300a, a total of four cylindrical propulsion units (for example, 90) rotatably disposed on, for example, both the front and rear sides of the vehicle 300a, and each cylindrical propulsion unit 90. A rotation mechanism (not shown) that rotates is configured. In FIG. 31, a wing 300b is further provided. Each cylindrical propulsion unit 90 is disposed in the vehicle 300a so as to be rotatable around its center of gravity, and can be rotated in the vertical direction or a horizontal direction parallel to the front-rear direction of the vehicle 300a by the rotation mechanism. It has become.

  With this configuration, if each tubular propulsion unit 90 is directed vertically and the wake is injected downward, stable vertical take-off and landing and hovering can be performed. If the wake is gently rotated in a substantially horizontal direction (arrow direction S in FIG. 31) parallel to the front-rear direction of the vehicle 300a and the wake is ejected to the rear of the vehicle, it is possible to shift to a horizontal flight in a stable state.

  In addition, since the fuselage or cylindrical propulsion unit described above is not only used as a flying device but also excellent in stability, it can also be used as an attitude control device for enhancing the stability of an airplane or the like. It is included in the scope of the above embodiment.

  In the above description, the airframe 90 is used as the cylindrical propulsion unit. However, the airframe 90 is not limited to the airframe 90, and may be the airframes 70, 80 or the like. Further, there is no particular limitation as long as the propulsion unit includes a cylindrical stabilizing blade or a radial stabilizing blade as a stabilizing blade.

Claims (33)

  A fuselage having a plurality of vertical partition plates assembled radially;
  A propeller disposed above the plurality of vertical partition plates;
With
  The diameter of the propeller is r ₀ 2πr of the moment that the propeller receives from the surrounding air ₀ The shape of the vertical partition plate is formed such that a double moment is applied to the plurality of vertical partition plates by the propeller wake.
  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,
  Δl represents the height of the i-th layer and the number of diameter units of the vertical partition plate. _i And m _i Then, the height Δl of each layer ₁ , Δl ₂ , ..., Δl _n And the number m of vertical partition plate diameter units ₁ , M ₂ , ..., m _n Is Δl ₁ m ₁ + Δl ₂ m ₂ + ... + Δl _n m _n ≒ 2πr ₀ A propeller device that is set to satisfy 2. The propeller device according to claim 1, wherein each of the layers further includes a cylindrical body that is covered around the plurality of vertical partition plates. With radial stabilizer or cylindrical stabilizer,
  The radial stabilizer blade and / or the cylindrical stabilizer blade is a distance n between the center of gravity of the attitude control device and the total wake wind pressure center point of the wind pressure center point of each stabilizer blade. _GC And the distance n between the center of gravity and the external wind pressure center point W _GW The posture control device is arranged such that the relationship between the position and the position is represented by Expression 19- (5).

A cylindrical stabilizer,
  One or more radial stabilizers disposed coaxially along the central axis of the cylindrical stabilizer;
With
  The radial stabilizer blade and / or the cylindrical stabilizer blade is a distance n between the center of gravity of the attitude control device and the total wake wind pressure center point of the wind pressure center point of each stabilizer blade. _GC And the distance n between the center of gravity and the external wind pressure center point W _GW The posture control device is arranged such that the relationship between the position and the position is represented by Expression 19- (5).

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 3 or 4, characterized by the above-mentioned. The attitude control device according to claim 3, wherein the wind flow generation device includes a propeller that generates a wind flow and a propeller drive unit. The attitude control device according to any one of claims 3 to 5, wherein an auxiliary stabilizer blade is disposed under the cylindrical stabilizer blade. In the case where the radial stabilizer is formed symmetrically about the central axis,
  The effective angle of the propeller is β _T And the number of the wings in the diameter unit of the i-th radial wing is m _i And the value obtained by dividing the length of the i-th radial stabilizer blade in the direction of the central axis by the diameter of the propeller is n _i Then, the posture control apparatus according to claim 3, wherein Expression 14- (1) is established.

An attitude control apparatus combining two of the same attitude control apparatuses according to any one of claims 3 to 8,
  The intake side open ends are directed upward, the exhaust side open ends are directed downward, and the intake side open ends are inclined at an inclination angle of 0 ° in the opposite direction to each other and spaced apart from each other. The two attitude control devices arranged; and
  A connecting member that interconnects the two attitude control devices;
An attitude control device comprising: A posture control device that is a combination of two or more posture control devices having the inclination angle larger than 0 ° of the posture control device according to claim 9, wherein the two or more posture control devices are viewed in plan view. A posture control device characterized by being interconnected so as to cross radially at the central axis of the. The attitude control device according to claim 9 or 10, wherein a distance n between the total gravity center G and the total wake wind pressure center point C _GC And the distance n between the total center of gravity G and the total external wind pressure center point W _GW Satisfies the expression 19- (5).

11. The attitude control device according to claim 10, wherein the number of the attitude control devices to be combined is an even number of 4 or more, and the total center of gravity, the total wake wind pressure center point, and the total external wind pressure center point coincide with each other. A posture control device characterized by having been made. The attitude control device according to any one of claims 3 to 12, further comprising a stabilizer blade outside the cylindrical stabilizer blade. An attitude control apparatus combining two or more of the attitude control apparatuses according to any one of claims 3 to 8 or claim 13,
  The two or more attitude control devices arranged such that their central axes are parallel to each other, their intake side is directed upward, their exhaust side is directed downward, and spaced apart from each other;
  A connecting member that interconnects the two or more attitude control devices;
An attitude control device comprising: 15. The attitude control device according to claim 14, wherein the total center of gravity G ₀ And the wake wind pressure center point C of each of the two or more attitude control devices _i The distance n in the central axis direction (i: the identification number of the two or more attitude control devices) _GCi , And its total center of gravity G ₀ And its total external wind pressure center point W ₀ And the distance n in the central axis direction _GW The posture control device characterized in that the relational expression satisfies the expression 21- (44) -3.

15. The attitude control device according to claim 14, wherein the total center of gravity G ₀ And its overall wake wind pressure central point C ₀ And an attitude control device characterized by matching each other. A propulsion power amplifying device using as a propulsion unit a combination of two or more of the attitude control devices according to claim 15 or claim 16,
  The two or more attitude control devices arranged substantially symmetrically so as to direct the propulsive force in the same rotational direction along a coaxial circumference;
  A connecting member that interconnects the two or more attitude control devices;
A propulsion power amplifying device comprising: The propulsion power amplifying device according to claim 17,
  Each of the two or more attitude control devices is characterized in that a straight line connecting the propulsive force action point and the center of the coaxial circumference and the central axis are interconnected so as to be orthogonal to each other. apparatus. 15. A posture control device that combines a plurality of the posture control devices according to claim 3, wherein the total thrust force action point, the total center of gravity, the total wake wind pressure center An attitude control device characterized in that a point and a central point of its total external wind pressure are arranged on its central axis. A flying device using as a propulsion unit a combination of two of the same attitude control devices according to any one of claims 3 to 8,
  The intake side opening ends are directed upward, the exhaust side opening ends are directed downward, and the intake side opening ends are opposed to each other at an inclination angle of 0 °, and spaced apart from each other. The two attitude control devices arranged; and
  A connecting member that interconnects the two attitude control devices;
A flight apparatus comprising: A flying device using the attitude control device according to any one of claims 3 to 12 as a propulsion unit,
  A flying device further comprising a stabilizing wing outside the cylindrical stabilizing wing. A flying device using as a propulsion unit a combination of two or more of the same attitude control devices according to any one of claims 3 to 8 or claim 13,
  The two or more attitude control devices arranged such that their central axes are parallel to each other, their intake side is directed upward, their exhaust side is directed downward, and spaced apart from each other;
  A connecting member that interconnects the two or more attitude control devices;
A flight apparatus comprising: 23. The flying device according to claim 22, wherein the overall center of gravity G ₀ And the wake wind pressure center point C of each of the two or more attitude control devices _i The distance n in the central axis direction (i: the identification number of the two or more attitude control devices) _GCi , And its total center of gravity G ₀ And its total external wind pressure center point W ₀ And the distance n in the central axis direction _GW And the relational expression satisfies the expression 21- (44) -3.

A flying device using the attitude control device according to claim 14 as a propulsion unit, the overall center of gravity G ₀ And its overall wake wind pressure central point C ₀ And a flight device characterized by matching each other. A flying device using as a propulsion unit a combination of a plurality of the attitude control devices according to any one of claims 3 to 12 or claim 14, wherein the total propulsive force action point, the total center of gravity, The flight device characterized in that the overall wake wind pressure center point and the overall external wind pressure center point are arranged on the center axis. The attitude control device according to any one of claims 3 to 16, wherein the propulsive force amplifying device according to claim 17 or 18 is used as a wind flow generating device. The flying device according to any one of claims 20 to 25, wherein the attitude control device according to claim 26 is used as a propulsion unit. A flying device according to any of claims 20 to 25 or claim 27,
  A flying device further comprising an auxiliary wing so as to cancel a moment around the total center of gravity caused by external wind pressure. The propulsion power amplifying device according to claim 17 or 18,
  The propulsion force amplifying apparatus according to claim 1, further comprising an auxiliary wing in each propulsion unit so as to cancel a moment around the total center of gravity of each propulsion unit due to an external wind flow pressure. The posture control device according to claim 19, wherein the posture control device satisfies Expression 21- (44) -3.

26. The flying device according to claim 25, wherein the flying device satisfies Formula 21- (44) -3.

The posture control device according to claim 30, wherein the central axes thereof are parallel to each other, the intake sides thereof are directed upward, the exhaust sides thereof are directed downward, and are spaced apart from each other. Each propulsion unit and a connecting member that interconnects the propulsion units, and each propulsion unit is set so that its center of gravity and the wake wind pressure center point coincide with each other, and each wake wind pressure It is installed so that it can rotate in the same direction around the center point or each center of gravity, and the overall center of gravity, overall wake wind pressure center point, overall external wind pressure center point, and overall external An attitude control device characterized by matching a wind current and a wind pressure center point. 32. Each of the propulsion units according to claim 31, wherein the propulsion units are arranged with their central axes parallel to each other, with their intake sides facing up and their exhaust sides facing down and spaced apart from each other. And a connecting member for interconnecting the propulsion units, and each propulsion unit is set so that the center of gravity and the wake wind pressure center point coincide with each other, and the wake wind pressure center point or It is installed so that it can rotate in the same direction around each center of gravity, and the overall center of gravity, overall wake wind pressure center point, overall external wind pressure center point, and overall external wind wind pressure center point of the entire flight device And a flight device characterized in that

Images