JP2008093204A - Co-axial helicopter - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a co-axial helicopter which is capable of moving in a horizontal direction with a large progressive thrust and excellent in flight stability when flying forward. <P>SOLUTION: The co-axial helicopter 10 comprises: an airframe body 11 in which upper and lower parts open and side part is all surrounded by a duct 21; and an upper part main rotor 22 and a lower part main rotor 23 which are supported co-axially along a main rotation axis 27 extended upward from the airframe body and rotate around the main rotation axis in the opposite direction to each other. The helicopter is equipped with: a vertical tail 41 extended backward from a duct 21a on the side of a tail; a pitching propeller 42 which is installed to an upper end of the vertical tail and is supported by a sub-rotation axis extended upward; and a tilted tail 44 installed on the side on both sides of the vertical tail. <P>COPYRIGHT: (C)2008,JPO&INPIT

Description

  The present invention relates to a counter-rotating rotary wing aircraft that is arranged coaxially along a rotation axis and flies with two upper and lower main rotors rotating in opposite directions, and more particularly in an environment affected by wind. The present invention relates to an unmanned small counter-rotating rotary wing aircraft that can stably fly.

Compared to fixed wing aircraft, small rotary wing aircraft have advantages such as not requiring a large area for take-off and landing, and capable of hovering (stop in the air), not only for radio control helicopters for hobbies It is used as an unmanned aerial vehicle (UAV (Unmanned Aerial Vehicle)) for the purpose of observation from above.
The basic structure of a small rotorcraft that has been used in the past is to cancel the torque in the horizontal direction (yaw direction) generated by the rotation of the main rotor and the main rotor provided at the top of the fuselage. A single rotor type helicopter having a vertical rotor for performing direction control is generally used.

  The single rotor type helicopter has a feature that the control of the main rotor and the tail vertical rotor (tail rotor) is relatively easy, but the engine is not always used for the generation of lift during flight. Part of the output must be supplied. In addition, the range of rotation of the vertical rotor is included up to the relatively low position of the aircraft, so if the aircraft loses its balance during takeoff and landing and the tail is lowered, the vertical rotor contacts the ground and ground objects. There is a risk of end.

  On the other hand, as one of the small rotor blades with different basic structures, the upper main rotor and the lower main rotor rotating in opposite directions are arranged on the same axis, so that strong lift is generated by the rotation of these two main rotors. In addition, a counter-rotating rotorcraft has been developed that simultaneously cancels torque in the yaw direction, thereby eliminating the need for a vertical rotor. The contra-rotating rotorcraft can reduce the rotor diameter for obtaining the same lift as compared with the single rotor type.

  On the other hand, since there is no vertical rotor in the counter-rotating rotary wing aircraft, the direction control of the nose must be performed by means other than the vertical rotor. Therefore, for example, a rotor blade pitch angle adjustment mechanism (swash plate) is provided for each of the upper and lower main rotors, and the pitch angle of the blades of the upper and lower two main rotors is simultaneously changed in the opposite direction to keep the total sum of lifts constant. It is disclosed that direction control is performed by breaking the balance of torque in the horizontal direction (see Patent Document 1).

  In addition, the indoor helicopter with counter rotating rotor blades is equipped with a stabilizer bar that is linked to the upper main rotor of the counter rotating rotor blades for the purpose of ensuring stable hovering performance and driving safety. A helicopter is disclosed (see Patent Document 2). That is, the upper main rotor attached to the mast (rotor rotation shaft) and the upper end of the mast are attached to the upper end of the mast so as to be tiltable with respect to the mast, and an appropriate crossing angle is provided via a link rod to the blade tilting mechanism of the upper main rotor And a stabilizer bar that is connected so as to be rotated and attached to rotate in conjunction with the upper main rotor is disclosed (see Patent Document 2).

FIG. 7 is a schematic diagram for explaining the posture stabilizing action by the stabilizer bar. When the airframe is hovering in a horizontal posture, as shown in FIG. 7A, the stabilizer bar 101 is rotated so that the rotation surface is horizontal due to the centrifugal force. The stabilizer bar 101 is provided with weights 102 at both ends of the bar so as to increase the moment of inertia. As a result, the rotation surface of the stabilizer bar 101 tends to maintain a horizontal level. If the aircraft tilts for some reason, as shown in FIG. 7 (b), the rotation surface of the upper main rotor 103 (and the lower main rotor 104) is opposite to the rotation surface of the stabilizer bar 101 which is intended to maintain the horizontal state. At this time, the link rod 105 is actuated to tilt the upper main rotor 103, so that the cyclic pitch for returning the stabilizer bar 101 and the upper main rotor 103 to be parallel is set to the upper main rotor 103. Input to the rotor. That is, without actively controlling by an external electric control signal, the centrifugal force acting on the stabilizer bar mechanically attempts to return the rotation surface of the upper main rotor to the upper main rotor horizontally. The effect is produced.
Therefore, when the airframe is tilted due to an external factor, the stabilizer bar 101 is kept horizontal, and the upper main rotor 103 tilted with respect to the stabilizer bar 101 functions as a correction rod to correct the attitude of the airframe horizontally. Thus, a restoring force that returns the aircraft to a horizontal posture is generated.

In addition, under the two upper and lower horizontal rotor blades, the remote control and unmanned small size that can be operated up and down by controlling the area adjustment means that can adjust the projected area that receives the downward flow generated by the two rotor blades. A flying object is disclosed (see Patent Document 3).
In the present invention, it is further disclosed that an axisymmetric cylinder is provided as a cover for covering the horizontal rotary blades (main rotor) with respect to the rotation shafts of the two horizontal rotary blades (main rotor). That is, by providing a cover that covers the side surface of a horizontal rotating wing on a flying object having a counter-rotating horizontal rotating wing, it is possible to operate safely without complicating the structure like a conventional helicopter. It is disclosed that it becomes possible.
JP-A-1-101297 Japanese Patent No. 3723820 JP 11-115896 A

  Small rotorcraft having a structure using counter-rotating rotor blades, although there is a potential demand in various fields, have not been used so far. The main reason for this is that it was difficult to perform stable and controllable flight in an environment affected by resistance such as wind. In general, the smaller the flying object, the stronger the influence of resistance due to wind or the like, and the more difficult it is to control to move to a desired position or to hover (stop in the air) to a certain position.

  In particular, in a small counter rotating rotor, the counter rotating rotor must play three roles: upward thrust against its own weight, thrust against resistance such as wind, and moment generation for posture control. However, due to the structure of the counter-rotating rotary blade, the thrust generated by the rotation of the counter-rotating rotary blade is much higher in the ratio of the thrust toward the upper side than the thrust in the horizontal direction. By controlling the pitch angle of the blade mounted on the counter rotating rotor blade and giving a so-called cycle pitch, it is possible to generate a thrust in the horizontal direction, but the thrust generated by the cycle pitch alone is sufficient for crosswind resistance. You can't overcome. As a result, the aircraft was swept away by the wind. Therefore, the small counter-rotating rotary wing aircraft is restricted to use indoors where the influence of wind is small.

In addition, since the conventional counter-rotating rotary wing aircraft does not fly forward with a large thrust, it has not been particularly intended to stabilize the aircraft in consideration of the attitude of the aircraft during forward flight.
For example, in Patent Document 2, stabilization is performed by a stabilizer bar. This is stabilization for returning the aircraft to a horizontal position when the aircraft is tilted, and aims to stabilize the aircraft posture during forward flight. Rather, it is intended to obtain optimal posture stability in the hovering state.

  Therefore, even if the present invention is a counter-rotating rotorcraft, it can move in the horizontal direction with a sufficiently large forward thrust as compared with the conventional ones, and even in an environment affected by the wind, it is swept away by the wind. It is an object of the present invention to provide a counter-rotating rotary wing aircraft that can fly in a desired direction and position while facing the wind, and that has excellent flight stability when flying forward.

  The contra-rotating rotary wing machine of the present invention, which has been made to solve the above-mentioned problems, includes a main body that is open at the top and bottom and is surrounded by ducts on the entire side, and a main rotation that extends upward from the main body. A counter-rotating rotor blade having an upper main rotor and a lower main rotor that are supported on the same axis along the axis and rotate in opposite directions around the main rotation axis, and directed backward from the rear duct A vertical tail, a pitching propeller attached to the upper end of the vertical tail and supported by an auxiliary rotating shaft extending upward, and inclined tails attached to the left and right side surfaces of the vertical tail. .

  Here, the duct that covers the side of the fuselage main body over the entire circumference is used as a counter rotating rotor when the upper main rotor and the lower main rotor (hereinafter, these two are also referred to as counter rotating rotor) rotate. On the other hand, it may be attached so that the duct does not contact. Specifically, when the lower main rotor of the counter rotating rotor or the upper and lower two main rotors are arranged so as to be housed in the duct, the inner diameter of the duct is configured to be larger than the rotation radius of the rotor. . Further, when the two upper and lower main rotors are arranged above the duct, they may be the same as the duct diameter or larger than the duct diameter.

  According to the present invention, the pitching propeller rotates to generate a pitching moment (head-down moment) that raises the tail and lowers the nose. As a result, the fuselage is inclined forward, and a part of the upward thrust (lift) by the upper and lower two main rotors (also referred to as counter rotating rotors) is a horizontal thrust. Due to this horizontal thrust, the aircraft flies forward with a large thrust. During forward flight, the vertical tail that extends rearward from the aft-side duct stabilizes the roll direction during forward flight and helps to fly straight.

  On the other hand, tilted tails are attached to the left and right side surfaces of the vertical tail, and the tilted tail is attached to a horizontal plane when the aircraft flies in a predetermined forward tilt posture. The predetermined forward leaning posture here is a forward leaning posture set as a reference posture during flight, and in principle, forward flight is performed in this forward leaning posture. By flying in such a way that the wing surface of the inclined tail wing is horizontal, the pitch direction of the aircraft is stabilized, and it helps to fly straight.

  According to the present invention, since a part of the large thrust generated by the counter rotating rotor blade is directly used as the thrust in the horizontal direction, the thrust force is sufficiently large as compared with the conventional counter rotating rotor aircraft. The aircraft structure can move in the horizontal direction, and the vertical and inclined tails stabilize the pitch and roll directions during forward flight. Even if there is, stable flight can be performed.

(Means and effects for solving other problems)
In the above invention, the mounting angle of the inclined tail is -30 degrees to -60 degrees with respect to the horizontal plane, and the rotational speed of the pitching propeller is set so that the inclined tail becomes horizontal when the aircraft flies. Also good.
By setting the forward leaning posture to -30 degrees to -60 degrees with respect to the horizontal plane, the upward thrust for flying and the horizontal thrust for moving forward can be distributed in a well-balanced manner, enabling stable forward flight. .

In the above invention, the inclined tail may be provided with an inclination angle adjusting mechanism for adjusting an angle with respect to a horizontal plane, and the angle of the inclined tail may be adjusted according to the inclination angle of the airframe.
According to the present invention, the tilt angle adjusting mechanism adjusts the angle according to the forward tilt angle of the airframe and controls the tilted tail in flight to be horizontal. Thereby, stable flight can be performed regardless of the forward tilt angle.

  Hereinafter, embodiments of the present invention will be described with reference to the drawings. The embodiment described below is merely an example, and modifications can be made without departing from the scope of the present invention.

(Airframe structure)
FIG. 1 is a perspective view showing an airframe structure of a small counter-rotating rotary wing aircraft according to an embodiment of the present invention, and FIG. 2 is a front view thereof (for convenience of explanation, a part of a duct is shown broken away). It is.
The small counter-rotating rotary wing machine 10 mainly includes a machine body 11 mounted with a control device and a drive mechanism and surrounded by a duct 21, an upper main rotor 22 and a lower main rotor. 23, a rotary blade mechanism 12 having a stabilizer bar 24, and a rear mechanism 13 having a vertical tail 41 and a pitching propeller 42 disposed behind the aft-side duct 21 a of the body body 11.

  The body body 11 is surrounded by a cylindrical duct 21 around the entire side surface, and the upper surface and the lower surface of the body body 11 are open. The duct 21 uses, for example, a thin-walled polystyrene foam in order to reduce the weight of the airframe as much as possible. The outer diameter of the duct 21 is approximately equal to the rotational radius of the two upper and lower main rotors, and the thrust ratio to the occupied area of the fuselage is as large as possible. The duct 21 rectifies the downward flow generated by the counter rotating rotor mechanism 12 to make it easy to control the upward thrust (lift). Plays a role in preventing contact. In addition to these, the duct 21 plays an important role in relation to the attitude control action by the stabilizer bar 24, which will be described later.

  A coaxial main rotating shaft 27 including an inner shaft 25 and an outer shaft 26 is rotatably supported by the frame 14 and a bearing (not shown) at the center of the body body 11 (on the central axis of the duct 21).

The inner shaft 25 is configured such that rotational power from a motor 28 mounted in the duct 21 is transmitted via a gear 29 on the lower end side thereof. A little below the upper end of the inner shaft 25, the pair of blades 22a of the upper main rotor 22 is pivotally supported via a blade tilting mechanism 22b. )) Is performed. A stabilizer bar 24 is pivotally supported at the upper end of the inner shaft 25 so as to be freely tiltable. Weights are attached to both ends of the stabilizer bar 24 so that the moment of inertia of the stabilizer bar 24 is increased.

The stabilizer bar 24 and the blade tilting mechanism 22 b of the upper main rotor 22 are connected by a link rod 31. The link rod 31 is for performing a posture control action, and is configured so that the stabilizer bar 24 and the upper main rotor 22 are rotated so as to maintain an appropriate crossing angle. The appropriate crossing angle here is determined in relation to the diameter of the main body 11, the length of the upper main rotor 22, the weight of the weight of the stabilizer bar 24, and the like, and is set to an optimum value by actual measurement. However, functionally, a restoring force is generated by inputting a cyclic pitch that restores the attitude of the aircraft horizontally when an external force is applied to the aircraft in the hovering state (stopped in the air) and the aircraft is tilted. This is a crossing angle.

Here, the attitude control action will be described. The stabilizer bar 24 and the upper main rotor 22 are configured to exert a posture stabilizing effect in a state where the airframe is inclined in the traveling direction by posture control by a combination of the pitching propeller 42 and the duct 21. That is, as shown in FIG. 5, during forward flight, the head-lowering moment M1 around the center of gravity G of the aircraft caused by the wind resistance N acting on the cylindrical surface of the duct 21 and the head-down moment M3 by the pitching propeller 42 The stabilizer bar 24, the link rod 31, and the upper main rotor 22 balance the action of the restoring moment M2 around the center of gravity G to return the airframe to a horizontal posture so that the aircraft can stably fly forward with the airframe tilted. It has become.

  The outer shaft 26 is configured such that rotational power from a motor 32 mounted in the duct 21 is transmitted via a gear 33 on the lower end side thereof. At the upper end of the outer shaft 26, a pair of blades 23a of the lower main rotor 23 is pivotally supported via a blade tilting mechanism 23b. The motor 32 is driven in a direction opposite to the upper main rotor 22 (reverse direction). )) Is performed.

  Further, in the machine body 11 (in the duct 21), two servo motors 34 and 35 for inputting a cyclic pitch for controlling the movement direction of the machine body are provided in the blade 23a of the lower main rotor 23. Yes. The cyclic pitch applied to the lower main rotor 23 is input via a rod (not shown) that is moved by these servo motors. The mechanical components of the rotary blade mechanism 12 including the main rotary shaft 27 (inner shaft 25, outer shaft 26), the upper main rotor 22, the lower main rotor 23, the stabilizer bar 24, the link rod 31 and the like are, for example, Hirobo's XRB aircraft (RC helicopter) can be used.

  In addition to the above, the airframe body 11 includes various sensors, specifically, an altimeter 37 that measures the airframe height by measuring atmospheric pressure, a GPS sensor 38a (GPS receiver) that obtains position information from GPS signals from satellites, and geomagnetic measurement. The direction sensor 38b for measuring the direction of the aircraft by the above, and the control device 39, the battery 40, and the radio control receiver 46 for receiving the steering signal are attached.

  A vertical tail 41 is fixed to the aft-side duct 21a of the body body 11 toward the rear. Further, a small pitching propeller 42 and a drive motor 43 that are rotated in parallel with the upper main rotor 22 and the lower main rotor 23 are attached to the upper end of the vertical tail 41 at a position that does not contact the upper main rotor 22 and the lower main rotor 23. The pitching propeller 42 tilts the fuselage main body 11 forward by giving a pitching moment (head-lowering moment) to lift the tail of the fuselage main body 11, and upward thrust by the upper main rotor 22 and the lower main rotor 23. A part of can be used as thrust for forward movement.

A pair of inclined tails 44 are attached to the left and right side surfaces of the vertical tail 41. The inclined tail blade 44 is operated so that the wing surface of the inclined tail blade 44 becomes substantially horizontal when the aircraft flies forward in a forward inclined state in which the nose side is lifted and the nose side is lowered by the operation of the pitching propeller 42. The tilt angle is set in Note that the forward leaning state of the aircraft during flight changes according to the rotation speed of the pitching propeller, so a reference forward flight speed is set in advance, and the inclination of the inclined tail is determined according to the rotation speed of the pitching propeller at that time. Set the mounting angle.
In the aircraft of the present embodiment, the flight at a forward flight speed of 2 m / s is set as a reference, and the forward tilt angle of the aircraft at this time is 45 degrees forward from the horizontal plane. Is attached at -45 degrees.
When the forward flight speed is slightly reduced, the forward tilt angle of the aircraft is made shallower, and when the forward flight speed is made slightly faster, the forward tilt angle of the aircraft is increased. Specifically, an appropriate angle is set between 30 degrees and 60 degrees. And an attachment angle is also set between -30 degrees--60 degrees according to this.

(Control system)
Next, a control system of the small counter-rotating rotary wing machine 10 will be described. FIG. 3 is a functional block diagram for explaining the configuration of the control system of the small contra-rotating rotary wing machine 10 and the input / output control signals of the respective parts constituting the control system. FIG. 4 is a diagram for explaining the relationship between measurement data and calculation data used for control.

  The control system is mechanically linked by the action of the stabilizer bar 24 and the link rod 31 of the rotary blade mechanism 12 in addition to the microcomputer control system A that performs guidance control by a control signal from the control device 39 mounted in the body 11. It has a mechanical control system B that controls the attitude of the fuselage by operation.

  As shown in FIG. 3, the microcomputer control system A of the small counter-rotating rotary wing machine 10 includes a control device 39 including a microcomputer (CPU) 39a for executing various operations and processes, and a memory 39b for storing necessary information. An altimeter 37, a GPS sensor 38a, an azimuth sensor 38b, and a radio control receiver 46 that receives a control signal when the operator manually controls the aircraft (ie, by operating the radio control transmitter 47). . On the other hand, the mechanical control system B includes a stabilizer bar 24, a link rod 31, and the upper main rotor 22, and the duct 21 is related to the mechanical control system B.

  Processing and functions executed by the microcomputer 39a and the memory 39b of the microcomputer control system A will be described for each functional block. The memory 39b includes a target point information storage unit 61. The microcomputer 39a includes a current point detection unit 51 and a horizontal displacement calculation unit 52. The current altitude detecting unit 53, the vertical displacement calculating unit 54, and the rotor control unit 55 are provided.

  The target point information storage unit 61 stores spatial coordinates (X, Y, Z) composed of longitude, latitude, and altitude as the spatial coordinates of the target point P to be reached by the aircraft. Of these, longitude (X coordinate) and latitude (Y coordinate) are horizontal components of spatial coordinates, and altitude (Z coordinate) is an altitude component. The spatial coordinates can be set by connecting a personal computer (not shown) and the control device 39 and performing an input operation from a keyboard.

The current location detection unit 51 detects the spatial coordinates (x, y, z) including the longitude, latitude, and altitude of the position Q where the aircraft is currently located (referred to as the current location Q) from the GPS signal received by the GPS sensor 38a. To do.
The horizontal displacement calculation unit 52 is (X, Y) which is the horizontal component of the coordinates (X, Y, Z) of the target point P and the horizontal component of the coordinates (x, y, z) of the current point Q (x , Y), a calculation is performed to calculate a horizontal target azimuth (θ) and a horizontal distance (L) to the target point to move the aircraft.

The current altitude detection unit 53 detects the current altitude (z) at which the aircraft is currently present from the altitude signal measured by the altimeter 37. Although the current altitude can be obtained from the GPS sensor 38a, it is more preferable to use altitude information from an altimeter 37 provided separately from the GPS sensor 38a from the viewpoint of accuracy and reliability.
The vertical displacement calculation unit 54 (Z) which is the altitude component of the coordinates (X, Y, Z) of the target point P and (z) which is the altitude component of the coordinates (x, y, z) of the current point Q Based on the above, the calculation for calculating the altitude difference (H) to the target point is performed.

  The rotor control unit 55 calculates the calculated horizontal target azimuth (θ), the horizontal distance to the target point (L), the altitude difference (H) to the target point, and the current nose direction obtained by the azimuth sensor. Based on (ψ), an altitude command, a yaw angle command, a horizontal position command (advance speed command) are generated, and signals sent from the altimeter 37, the GPS sensor 38a, and the direction sensor 38b are detected every moment, By performing feedback control to update each command, the rotational speed of the upper main rotor 22, the rotational speed of the lower main rotor 23, the rotational speed of the pitching propeller 42, and the cyclic pitch to be input to the lower main rotor 23 (for vertical flaps) By appropriately controlling the cyclic pitch of the aircraft, it is possible to control the height of the aircraft, control the direction of aircraft movement, If not, control the aircraft's forward speed as well as the horizontal position.

The above is the guidance control by the microcomputer control system A, but in the airframe equipped with the mechanical control system B by the stabilizer bar 24, the control for the turning of the airframe is additionally executed.
That is, the pitching propeller 42 is operated for forward flight, and the duct 21 receives wind resistance from the front, so that the airframe changes from the horizontal state to the forward inclined state, so that the mechanical control system B operates, and FIG. As described above, the stabilizer bar 24 inputs a cyclic pitch to the upper main rotor 22 via the link rod 31. As a result, a restoring moment in the direction of returning the attitude of the aircraft from the tilted state to the horizontal state works separately. As a result, the action of tilting the aircraft by the microcomputer control system A competes with the action of restoring the horizontal state by the mechanical control system B, and these are balanced to slightly tilt the aircraft. Forward flight will be performed in the state.

By the way, the relationship between the stabilizer bar 24, the link rod 31, and the upper main rotor 22 constituting the mechanical control system B is, as described above, a restoring moment is generated so as to return it to a horizontal state when the airframe is tilted. In this way, the cyclic pitch is input, and is designed to be kept horizontal when in a hovering state (aerial stop state). For this reason, when the microcomputer control system A and the mechanical control system B compete with each other and the aircraft flies forward while maintaining a balance while the aircraft is tilted, the cyclic pitch for ensuring a stable forward flight is not necessarily the upper main rotor 22. It is not designed to be given to the other, but rather an incomplete cyclic pitch has been entered. Specifically, the cyclic pitch at which the stabilizer bar 24 and the link rod 31 are input to the upper main rotor 22 during forward flight in an inclined state is an effect of turning the aircraft in one direction as the thrust of forward flight increases. Will increase. This turning direction is determined to be either a right turn or a left turn depending on conditions such as the rotation direction of the upper main rotor 22 to which the stabilizer bar 24 is connected by the link rod 31 and the attachment method of the link rod 31. However, it is assumed here that the vehicle turns left for convenience.
Therefore, in addition to the guidance control described above, the rotor control unit 55 of the microcomputer control system A cyclically inputs to the lower main rotor 23 in order to cancel the left turn that occurs when the aircraft flies forward with the aircraft tilted. The rolling moment is controlled by the pitch (cyclic pitch for the lateral flap), and the yawing moment is controlled by generating the torque in the yaw direction by changing the rotation speed of the upper main rotor and the lower main rotor. By doing so, control for giving a right turn that balances with the left turn is added.

(Flying operation)
Next, the flight operation of the small counter rotating rotor aircraft 10 will be described.
When the coordinates (X, Y, Z) of the target point P are input in advance and the control is started, the upper main rotor 22 is rotated by the altitude command, the yaw angle command, and the horizontal position command which are generated every moment by the control device 39. By controlling the speed, the rotation speed of the lower main rotor 23, the rotation speed of the pitching propeller 42, and the cyclic pitch (cyclic pitch for the vertical and horizontal flaps) input to the lower main rotor 23, the direction of the aircraft is controlled. Fly toward the target point P. At this time, the aircraft flies forward while tilting forward about 45 degrees. As a result, the inclined tail wing flies while being in a substantially horizontal state.
At this time, the presence of the inclined tail 44 and the vertical tail 41 improves the stability of the pitch direction and the roll direction when the aircraft flies forward.

(Modified Example)
In the above embodiment, the inclined tail is fixed to the vertical tail. However, a sensor (gyro sensor) for measuring the inclination angle of the aircraft is provided in the aircraft body 11, and the angle of the inclined tail is determined according to the forward inclination angle of the aircraft. May be provided with an inclination angle adjusting mechanism that always becomes parallel to the horizontal plane. In this way, the angle of the inclined tail can be maintained at an optimum value according to the forward tilt attitude of the aircraft, and stable forward flight can be realized regardless of the forward tilt angle of the aircraft. become.

  The present invention can be used for an unmanned counter rotating rotary wing machine.

1 is a perspective view of a counter-rotating rotary wing aircraft according to an embodiment of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS The front view which shows the structure of the contra-rotating rotary blade machine which is one Embodiment of this invention (a part of duct is broken and it has shown the inside). The block block diagram which shows the structure of a control system, and the functional block diagram for demonstrating the control signal input / output by each part which comprises a control system. The figure explaining the relationship of the measurement data and calculation data required for control. The figure explaining the attitude | position control effect | action by a stabilizer bar and a duct. The figure explaining the attitude | position control effect | action by a stabilizer bar.

Explanation of symbols

10: Small counter rotating rotor blade 11: Airframe body 12: Rotor blade mechanism 13: Rear mechanism 21: Duct 22: Upper main rotor 23: Lower main rotor 24: Stabilizer bar 25: Inner shaft 26: Outer shaft 27: Main Rotating shaft 31: Link rod 37: Altimeter 38a: GPS sensor 38b: Direction sensor 39: Control device 41: Vertical tail 42: Pitching propeller 44: Inclined tail 51: Current position detection unit 52: Horizontal displacement calculation unit 53: Current altitude detection Unit 54: Vertical displacement calculation unit 55: Rotor control unit 61: Target point information storage unit

Claims (3)

The airframe main body whose upper and lower sides are open and the sides are surrounded by ducts, and the main rotational axis extending upward from the airframe main body are supported on the same axis, and around the main rotational axis in opposite directions. A counter-rotating rotary wing machine equipped with a rotating upper main rotor and a lower main rotor,
A vertical tail extending rearward from the aft-side duct;
A pitching propeller attached to the upper end of the vertical tail and supported by a secondary rotating shaft extending upward;
A counter-rotating rotary wing machine comprising an inclined tail mounted on the left and right sides of a vertical tail. The mounting angle of the inclined tail is -30 degrees to -60 degrees with respect to a horizontal plane, and the rotational speed of the pitching propeller is set so that the inclined tail becomes horizontal when the aircraft flies. The contra-rotating rotary wing machine according to 1. The counter-rotating rotary wing machine according to claim 1, wherein the inclined tail wing includes an inclination angle adjusting mechanism that adjusts an angle with respect to a horizontal plane, and the angle of the inclined tail wing is adjusted according to the inclination angle of the fuselage. .

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