JP2008094277A - Double reversal rotation impeller machine - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a double reversal rotation impeller machine capable of being moved in a horizontal direction by large advancement thrust, and easily, accurately horizontally moved and hovered in a desired direction and to a position even under an environment receiving influence of wind. <P>SOLUTION: The double reversal rotation impeller machine is provided with a machine body 21; an upper main rotor 22 and a lower main rotor 23; a stabilizer bar 24 supported inclinably relative to a main rotation shaft; and a link rod 31 for connecting the stabilizer bar and the upper main rotor. The machine is provided with a pitching propeller 42 mounted to a neighborhood to a machine tail side duct and supported on an upwardly extending sub-rotation shaft; and a rotor control part for performing linear advancement flight by balancing left and right turning forces by control of rolling moment by cyclic pitch given to a blade of the lower main rotor relative to the machine body being going to turn in any one direction of left and right directions by cyclic pitch of the upper main rotor 22 trying to return forward attitude to horizontal attitude. <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 forward with a large thrust.

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 order to improve the attitude stability of the counter rotating rotor aircraft even when affected by wind, a method of connecting the upper main rotor and the stabilizer bar with a link rod as described in Patent Document 2 is as follows. It is an effective countermeasure method.
However, even if the posture stability can be improved, it is difficult to fly with a large thrust without being swept away by the wind.

  In addition, by connecting the stabilizer bar to the upper main rotor, the posture stability can be improved and the posture can be kept horizontal in a hovering state or the like, while the aircraft is intentionally inclined (tilted forward). At the same time, the stabilizer bar is attached, and in addition to the restoring action to return to the horizontal, the action of turning to either the left or right works, so it is difficult to fly straight in the desired direction when performing forward flight The situation that became.

Therefore, even if the present invention is a counter-rotating rotary wing aircraft, it can move in the horizontal direction with a sufficiently large forward thrust compared to the conventional aircraft, and even in an environment affected by wind, An object of the present invention is to provide a small contra-rotating rotorcraft capable of flying in a desired direction and position while competing with the above.
It is another object of the present invention to provide a small counter-rotating rotary wing aircraft capable of solving the problem relating to the flight direction caused by the presence of a stabilizer bar during forward flight.

  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. 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 a rod-like body with weights for increasing the moment of inertia mounted at both ends, A stabilizer bar whose center is supported to be tiltable with respect to the main rotation shaft at the upper end of the main rotation shaft, and the stabilizer bar rotates integrally with the blade while maintaining an appropriate crossing angle with respect to the blade of the upper main rotor The stabilizer rod and the link rod that connects the upper main rotor, and the stabilizer that keeps it horizontal by the moment of inertia during the flight. When the rotation surface of the upper main rotor is inclined with respect to the rotation surface of-, the restoring force that attempts to return the rotation surface of the upper main rotor horizontally by the cyclic pitch that the link rod inputs to the blades of the upper main rotor In the counter-rotating rotary wing aircraft, the pitching propeller is mounted near the duct on the tail side and supported by the auxiliary rotating shaft that extends upward, and the heading-down moment around the center of gravity of the aircraft is reduced When the aircraft is flying in a forward leaning posture, the lower main rotor is used to turn the aircraft in one of the left and right directions by the cyclic pitch input of the upper main rotor that attempts to return the forward leaning posture to the horizontal posture. Control of the rolling moment by the cyclic pitch input to the blade of the upper main rotor By controlling the yawing moment by changing the rotational speed with the lower main rotor and generating torque in the yaw direction, the turning force for turning the aircraft in the direction opposite to the one direction is given to balance the left and right turning force And a rotor control unit for performing a straight flight.

  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.

  The auxiliary rotating shaft that supports the pitching propeller is supported so as to face upward at a position in the vicinity of the aft-side duct, but may be attached so as to be directly supported by the duct, or the aft-side duct. A tail wing (fixed wing) heading from the rear to the rear may be formed and supported on the tail wing such that the sub-rotation shaft faces upward.

According to the present invention, the aircraft body is levitated by rotating the two upper and lower main rotors (double counter-rotating rotor) to generate upward thrust. In this state, by further rotating the pitching propeller, a pitching moment (head-lowering moment) for lowering the nose and raising the tail is generated, and the aircraft tilts forward. As a result, a part of the thrust by the upper and lower two main rotors can be used as the thrust in the horizontal direction, which is larger than the horizontal thrust generated by the conventional cyclic pitch (for example, thrust ratio 1: 10) can be used.
By the way, when the machine body is tilted forward, the main rotation axis is tilted forward, so that the rotation surfaces of the upper and lower two main rotors are inclined with respect to the horizontal plane. On the other hand, the stabilizer bar which is supported on the main rotating shaft so as to be tiltable and has a large moment of inertia tends to maintain its rotating surface in a horizontal state. As a result, the rotation surface of the stabilizer bar and the rotation surface of the upper main rotor are not parallel. At this time, by adjusting the cyclic pitch that the link rod inputs to the blades of the upper main rotor, a restoring force is generated to return the rotating surface of the upper main rotor to the horizontal. As a result, the aircraft flies in a state where the force to tilt forward by the rotation of the pitching propeller and the force to return the forward tilt by the upper main rotor, link rod, and stabilizer bar to the horizontal are balanced.
At this time, as a side effect of the cyclic pitch input in which the upper main rotor attempts to return the forward leaning posture to the horizontal posture, a force is generated in the aircraft to turn in either the left or right direction, and the aircraft turns left or right. A rotating phenomenon occurs. This is because the stabilization effect by the stabilizer bar is adjusted by attaching the link rod so that the cyclic pitch is input so that the aircraft can maintain a horizontal state when hovering. This is due to the fact that it is not premised on.
Therefore, the rotor control unit performs rolling control based on the cyclic pitch (lateral cyclic pitch) input to the blades of the lower main rotor, and changes the rotational speeds of the upper main rotor and the lower main rotor to generate torque in the yaw direction. By controlling the yawing moment that is generated, the left and right turning forces are balanced and a straight flight is performed.

  According to the present invention, even with a counter-rotating rotary wing aircraft, a part of the thrust of the main rotor is used as the thrust for the horizontal movement, so that the forward thrust is sufficiently large compared to the conventional aircraft. It can move in the horizontal direction, and even in an environment affected by the wind, it can fly against the wind with a large thrust. According to the present invention, during the forward flight, the aircraft will fly in a forward leaning posture due to the rotation of the pitching propeller. At that time, the posture stabilizing function by the stabilizer bar, the link rod, and the upper main rotor is achieved. As a side effect, turning force in one of the left and right directions is generated, and this is controlled by the cyclic moment (lateral cyclic pitch) that the rotor controller inputs to the lower main rotor, and the upper main rotor and lower main It is possible to fly straight by balancing by controlling the yawing moment by changing the rotational speed with the rotor.

(Means and effects for solving other problems)
In the above invention, the rotor control unit generates the torque in the yaw direction by controlling the rolling moment by the cyclic pitch input to the blades of the lower main rotor and changing the rotational speeds of the upper main rotor and the lower main rotor. The magnitude of the reverse turning force given by controlling the yawing moment may be proportionally controlled (P control) according to the rotational speed of the pitching propeller.
The aircraft has a forward tilt angle controlled according to the rotational speed of the pitching propeller, and the forward speed changes according to the forward tilt angle. Since the magnitude of the turn is proportional to the forward speed, by controlling the magnitude of the reverse turning force to be proportional to the rotational speed of the pitching propeller, it is possible to fly straight without depending on the forward speed. Is possible.

Further, the contra-rotating rotary wing machine of the present invention made from another point of view includes an airframe body whose upper and lower sides are open and whose side is surrounded by a duct over the entire circumference, and a main rotating shaft extending upward from the airframe body. The upper main rotor and the lower main rotor that are supported on the same axis along the main axis of rotation and rotate in opposite directions around the main rotation axis, and the rod-shaped body with weights to increase the moment of inertia are attached to both ends. A stabilizer bar whose center is supported to be tiltable with respect to the main rotation shaft at the upper end of the main rotation shaft, and the stabilizer bar rotates integrally with the blade while maintaining an appropriate crossing angle with respect to the blade of the upper main rotor. A link rod that connects the stabilizer bar and the upper main rotor so that the stabilizer bar is kept horizontal by the moment of inertia during flight. When the rotating surface of the upper main rotor is inclined with respect to the rolling surface, a restoring force is generated to return the rotating surface of the upper main rotor to the horizontal by the cyclic pitch that the link rod inputs to the blade of the upper main rotor. In the contra-rotating rotary wing machine, the pitching propeller that is attached to the vicinity of the aft-side duct and supported by the auxiliary rotating shaft that extends upward, and the target point information storage unit that stores the spatial coordinates of the target point A current point detection unit that detects at least a horizontal component of the spatial coordinates of the current location by a GPS signal from a GPS receiver attached to the aircraft, a horizontal component of the spatial coordinates of the target location, and a horizontal component of the spatial coordinates of the current location A horizontal displacement calculation unit for calculating a horizontal target azimuth and a horizontal distance to the target point to move the aircraft based on And a yaw angle command for controlling the current heading of the nose to the horizontal target heading by changing the rotation speed of the upper main rotor and the lower main rotor. Horizontal position command to control the horizontal position of the aircraft by controlling the thrust in the horizontal direction by the upper main rotor and the lower main rotor by changing the rotation speed of the pitching propeller in proportion to the horizontal distance to the target point In addition, the rotor control unit is controlled by the cyclic pitch input of the upper main rotor that attempts to return the forward leaning posture to the horizontal posture when the aircraft is flying in the forward leaning posture by the pitching propeller. Input to the blade of the lower main rotor for the aircraft trying to turn in one direction The rolling moment is controlled by the cyclic pitch, and the yawing moment is controlled by changing the rotational speed of the upper main rotor and the lower main rotor to generate the yaw direction torque. Thus, a turning force for turning in the opposite direction is applied to balance the turning force on the left and right so that a straight flight can be performed.

  According to the present invention, the spatial coordinates of the target point are stored in advance in the target point information storage unit. When flying the aircraft with the spatial coordinates of the target point memorized, the GPS receiver receives the GPS signal sent from the artificial satellite during the flight and detects the space coordinates indicating the current location of the aircraft every moment. To do. The horizontal displacement calculation unit is based on the horizontal component of the spatial coordinates of the stored target point and the detected horizontal component of the spatial coordinate of the current point, and the horizontal target azimuth and the target point to be moved are determined. Calculate the horizontal distance. The direction sensor detects the current direction of the nose.

  Then, the rotor control unit first generates a yaw angle command for making the nose direction coincide with the horizontal target direction. The yaw angle command generates a torque in the yaw direction by changing the relative speed between the forward rotation speed of the upper main rotor and the reverse rotation speed of the lower main rotor, and rotates the nose direction. Take control. In the yaw angle control, the relative rotational speeds of the upper main rotor and the lower main rotor are maintained while keeping the sum of upward thrust (lift) by the upper and lower two main rotors (double reversing rotor) unchanged. If it is changed, the current altitude will be maintained, and conversely, if the relative rotational speed is changed while changing the sum of upward thrust (lift), the aircraft will be raised and lowered Therefore, along with the yaw angle control, altitude control is performed as necessary.

  Following the yaw angle control, the rotor control unit generates a horizontal position command for controlling the horizontal position of the aircraft. The horizontal position command performs control to generate a pitching moment (head-down moment) that raises the tail and lowers the nose by increasing the rotational speed of the pitching propeller. As a result, the fuselage is inclined forward, and a part of the upward thrust (lift) by the two upper and lower main rotors (double reversing rotors) is a horizontal thrust. Since this horizontal thrust changes depending on the forward tilt angle of the aircraft, control is performed to change the rotational speed of the pitching propeller in proportion to the horizontal distance to the target point. As a result, when the aircraft is swept away due to the influence of strong winds, it tries to return to the target point with a strong horizontal thrust. The horizontal thrust generated in the main rotor is much larger than the thrust generated by inputting a cyclic pitch to the blades of the main rotor (for example, the thrust ratio is 10: 1). Therefore, even in an environment affected by wind, the aircraft can be accurately brought close to the target point with a thrust according to the environment. In addition, when there is no influence of a wind, a forward speed can be enlarged compared with the past.

However, when the fuselage is tilted forward, the main rotation axis is tilted forward, so that the rotation surfaces of the upper and lower two main rotors are inclined with respect to the horizontal plane. On the other hand, the stabilizer bar which is supported on the main rotating shaft so as to be tiltable and has a large moment of inertia tends to maintain its rotating surface in a horizontal state. As a result, the rotation surface of the stabilizer bar and the rotation surface of the upper main rotor are not parallel. At this time, by adjusting the cyclic pitch that the link rod inputs to the blades of the upper main rotor, a restoring force is generated to return the rotating surface of the upper main rotor to the horizontal. As a result, the aircraft flies in a state where the force to tilt forward by the rotation of the pitching propeller and the force to return the forward tilt by the upper main rotor, link rod, and stabilizer bar to the horizontal are balanced.
At this time, as a side effect of the cyclic pitch input in which the upper main rotor attempts to return the forward leaning posture to the horizontal posture, a force is generated in the aircraft to turn in either the left or right direction, and the aircraft turns left or right. A rotating phenomenon occurs.
Therefore, in addition to the above control, the rotor control unit controls the rolling moment by the cyclic pitch (lateral cyclic pitch) input to the blades of the lower main rotor, and the rotational speed of the upper main rotor and the lower main rotor. By controlling the yawing moment by changing and generating the torque in the yaw direction, the left and right turning forces are balanced and a straight flight is performed.

  According to the present invention, during forward flight, the aircraft will fly in a forward leaning posture due to the rotation of the pitching propeller. At that time, as a side effect of the posture stabilization function by the stabilizer bar, the link rod, and the upper main rotor Rotating force in one of the left and right directions is generated, and this is controlled by the cyclic moment (lateral cyclic pitch) that the rotor control unit inputs to the lower main rotor, and the upper main rotor and lower main By making a balance by controlling the yawing moment by changing the rotational speed with the rotor and generating a torque in the yaw direction, it is possible to fly straight.

Further, in the above invention, a current altitude detecting unit that detects an altitude component of a spatial coordinate at a current location based on an altitude signal from an altimeter attached to the airframe; And a vertical displacement calculator that calculates an altitude difference based on the above, and the rotor controller changes the rotational speed of the upper main rotor and the lower main rotor in addition to the generation of the yaw angle command and the horizontal position command. Thus, an altitude command for controlling the current altitude of the aircraft to match the target altitude may be generated.
According to the present invention, not only horizontal movement but also vertical direction can be moved to a preset desired altitude, and stable hovering can be performed in a state controlled to the desired altitude.

  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. 6, during forward flight, a head-lowering moment M1 around the center of gravity G of the airframe caused by the wind resistance N acting on the cylindrical surface of the duct 21, and a 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 44 has an inclination angle so that the blade surface of the inclined tail 44 becomes substantially horizontal when the aircraft moves forward with the nose side raised and the nose side lowered. Specifically, for example, when the airframe body 11 is designed on the basis of moving forward in a state where it is inclined 45 degrees forward from the horizontal plane, it is attached so that the inclination angle becomes −45 degrees. The presence of the inclined tail 44 and the vertical tail 41 improves the stability in the pitch direction and the roll direction when the aircraft flies forward.
A sensor for measuring the tilt angle of the fuselage may be provided in the fuselage main body 11, and a tilt angle adjusting mechanism may be provided so that the angle of the tilt tail is always horizontal according to the tilt angle of the fuselage. In this way, the angle of the inclined tail can be maintained at an optimum value according to the inclination angle of the aircraft, so that more stable forward flight can be realized regardless of the inclination angle of the aircraft. Become.

(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 control by a control signal from the control device 39 mounted in the body 11. Thus, a mechanical control system B that performs control is provided. These two controls are performed independently. Mainly, the microcomputer control system A performs guidance control of the aircraft, and the mechanical control system B performs attitude control of the aircraft. May act to supplement or act as a complement.

  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) composed of 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 in which the mechanical control system B by the stabilizer bar 24 operates, 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 inputted, and it is designed so that the level is optimally maintained in the hovering state (air suspension 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. In other words, an incomplete cyclic pitch is input. Specifically, the cyclic pitch that the stabilizer bar 24 and the link rod 31 input to the upper main rotor 22 during forward flight in an inclined state has the effect of turning the aircraft in one direction as the forward flight thrust increases. It 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.

  Next, a specific control operation by the microcomputer control system A will be described. FIG. 5 is a diagram for explaining feedback control for airframe guidance executed by an altitude command, a yaw angle command, and a horizontal position command (forward speed command) generated by the rotor control unit 55.

(Advanced control)
Based on (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 When the altitude difference (H) is calculated, the rotational speeds of the upper main rotor 22 and the lower main rotor are controlled so that a vertical thrust (thrust) corresponding to the altitude difference (H) up to the target altitude is generated. Generate advanced commands. At this time, the rotational speed of the upper main rotor 22 and the rotational speed of the lower main rotor 23 are balanced so that no rotational force (torque) in the yaw direction is generated. The altitude command performs feedback control by PD control using the altitude difference (H) calculated based on the latest altitude signal sent from the altimeter 37 of the aircraft every moment. As a result, the target altitude (Z) is approached mainly by the P control component, and the responsiveness to rapid fluctuations is further enhanced by the D control component.

(Yaw angle control)
Based on (X, Y) which is the horizontal component of the coordinates (X, Y, Z) of the target point P and (x, y) which is the horizontal component of the coordinates (x, y, z) of the current point Q. When the horizontal target azimuth (θ) to be moved is calculated, the yaw angle command for controlling the yaw angle so that the nose direction coincides with the calculated horizontal target azimuth (θ). Is generated. That is, the current nose direction (ψ) is detected by the direction sensor 38b, the direction difference (φ) with respect to the target direction (θ) in the horizontal direction is calculated, and the upper main so that the direction difference (φ) approaches zero. A command for generating a rotational force (torque) in the yaw direction is generated by controlling the rotational speed of the rotor 22 and the rotational speed of the lower main rotor 23. At this time, the rotational speed of the upper main rotor 22 and the rotational speed of the lower main rotor are balanced so that the thrust (thrust) in the vertical direction does not change. The yaw angle command performs feedback control by PD control using the heading difference (φ) calculated based on the latest heading signal (ψ) sent from the heading sensor 38b every moment. Do.

(Horizontal position control)
Following the yaw angle control, the horizontal component (X, Y) 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) and the horizontal distance (L) is calculated, a horizontal position command (forward speed command) is generated to control the horizontal position so that the horizontal distance (L) approaches zero. This horizontal position command (forward speed command) includes the rotational speed of the pitching propeller 42 and the cyclic pitch for the vertical flap of the lower main rotor 23 (cyclic pitch control for generating forward flight thrust, which has been conventionally performed). And the cyclic pitch for the lateral flap of the lower main rotor 23 are controlled.

  That is, the horizontal position command (advance speed command) is controlled by controlling the rotational angle of the pitching propeller 42 and controlling the tilt angle of the airframe, so that the upper main rotor 22 and the lower main rotor 23 (double counter rotor) are controlled. ) Is used as a horizontal thrust, and the thrust is controlled. Second, the cyclic pitch input of the vertical flaps of the lower main rotor 23 is controlled to control the thrust in the horizontal direction. These two thrusts are considerably smaller than the former (for example, the thrust ratio is 10: 1), but by adding them together, the two thrusts can be advanced with as much thrust as possible.

Thirdly, the horizontal position command (advance speed command) inputs a cyclic pitch for the lateral flap of the lower main rotor 23 to control the rolling moment, and at the same time, the upper main rotor and the lower main rotor The yawing moment is controlled by changing the rotation speed and generating a torque in the yaw direction. As described above, the yaw moment is controlled by suppressing the phenomenon that the aircraft turns counterclockwise due to the influence of the attitude control by the stabilizer bar 24. The control which gives the cyclic pitch input of the lateral flap for making it move forward is performed.
The left-turning phenomenon tends to increase as the pitching propeller speed increases and the aircraft tilt angle increases, that is, the horizontal position command (advance speed command) increases. The cyclic pitch input for the lateral flap of the lower main rotor 23 and the rotational speed change for the upper main rotor 22 and the lower main rotor 23 for controlling the yawing moment are also controlled according to the horizontal position command. To do.

  The horizontal position command performs feedback control by P control (proportional control) using the latest horizontal distance (L) calculated based on the latest position information of the current location Q from the GPS sensor 38a. By performing feedback control by P control, the thrust in the horizontal direction can be increased as the distance from the target point increases. As a result, the stronger the wind, the more the aircraft will be washed away and the horizontal distance from the destination point tends to be longer, but since thrust is given in proportion to the horizontal distance, by increasing the thrust even when the wind is strong, You can reach the destination. On the other hand, when there is no wind, the vehicle is controlled to approach the destination at a forward speed proportional to the horizontal distance, so that the destination can be reached quickly and easily.

(General control)
The microcomputer control system A simultaneously performs aircraft guidance (including control for the left turn phenomenon) by altitude command, yaw angle command, and horizontal position command, but in addition to this, the operator can control the navigation command from the radio control transmitter 42. As shown in FIG. 5, when the radio control receiver 41 receives the steering signal, the control in which the steering command is added to the altitude command, the yaw angle command, and the horizontal position command is performed. Become. In this way, manual control can be performed in parallel, so that even if there is a sudden change, it can be manually handled.

  When performing manual control, by controlling the rotation speed of the pitching propeller, the tilt angle of the fuselage changes, which can change the thrust in the horizontal direction and control the forward speed, but depending on the forward speed As for the left turn phenomenon, the magnitude of the turn also changes. Therefore, the magnitude of the rightward turning force given by the lateral cyclic pitch input of the lower main rotor and the rotational speed change with respect to the upper main rotor 22 and the lower main rotor 23 is determined. The proportional control is performed in accordance with the rotational speed of the pitching propeller. Thereby, no matter how the forward speed changes, an appropriate right turning force can always be applied, and the right and left turning forces can be balanced and a straight flight can be realized.

  In the present embodiment, the three-dimensional target point P is determined in order to control the horizontal direction and the altitude direction. For example, when controlling only the horizontal direction, such as when flying at a constant altitude, the altitude direction control is performed. May not be performed.

  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 feedback control for the body guidance in the microcomputer control system of the counter rotating rotorcraft which is one embodiment of the present invention. 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 (4)

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. Rotating upper main rotor and lower main rotor;
A stabilizer bar comprising a rod-like body with weights for increasing the moment of inertia attached at both ends, the center of the rod-like body being supported tiltably with respect to the main rotation shaft at the upper end of the main rotation shaft;
A link rod that connects the stabilizer bar and the upper main rotor so that the stabilizer bar can rotate integrally with the blade of the upper main rotor while maintaining an appropriate crossing angle,
When the rotating surface of the upper main rotor is inclined with respect to the rotating surface of the stabilizer bar that is to be kept horizontal by the moment of inertia during flight, the upper main rotor is driven by the cyclic pitch that the link rod inputs to the blades of the upper main rotor. In the counter-rotating rotary blade that is designed to generate a restoring force that attempts to return the rotor's rotating surface to a horizontal position,
A pitching propeller attached to the vicinity of the aft-side duct and supported by a sub-rotating shaft extending upward,
When a pitching propeller is used to give a head-lowering moment around the center of gravity of the aircraft to fly the aircraft in a forward leaning posture, the upper main rotor that attempts to return the forward leaning posture to a horizontal posture is input in either the left or right direction. Controlling the rolling moment by the cyclic pitch input to the blade of the lower main rotor and generating the torque in the yaw direction by changing the rotational speed of the upper main rotor and the lower main rotor for the aircraft to turn And a rotor control unit that provides a turning force for turning the airframe in the direction opposite to the one direction by balancing the yawing moment by means of balancing the left and right turning forces to perform straight flight. Heavy reversal rotorcraft. The rolling moment caused by the cyclic pitch input to the blades of the lower main rotor by the rotor controller and the magnitude of the reverse turning force given by changing the rotational speed of the upper main rotor and the lower main rotor are determined by the pitching propeller. The counter rotating rotorcraft according to claim 1, wherein proportionally controlled (P control) is performed in accordance with the rotational speed of the rotor. 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. Rotating upper main rotor and lower main rotor;
A stabilizer bar comprising a rod-like body with weights for increasing the moment of inertia attached at both ends, the center of the rod-like body being supported tiltably with respect to the main rotation shaft at the upper end of the main rotation shaft;
A link rod that connects the stabilizer bar and the upper main rotor so that the stabilizer bar can rotate integrally with the blade of the upper main rotor while maintaining an appropriate crossing angle,
When the rotating surface of the upper main rotor is inclined with respect to the rotating surface of the stabilizer bar that is to be kept horizontal by the moment of inertia during flight, the upper main rotor is driven by the cyclic pitch that the link rod inputs to the blades of the upper main rotor. In the counter-rotating rotary blade that is designed to generate a restoring force that attempts to return the rotor's rotating surface to a horizontal position,
A pitching propeller attached to the vicinity of the aft-side duct and supported by a sub-rotating shaft extending upward,
A target point information storage unit for storing the spatial coordinates of the target point;
A current point detection unit that detects at least a horizontal component of the spatial coordinates of the current point by a GPS signal from a GPS receiver attached to the aircraft;
A horizontal displacement calculation unit that calculates a horizontal target azimuth and a horizontal distance to the target point to move the aircraft based on a horizontal component of the spatial coordinate of the target point and a horizontal component of the spatial coordinate of the current point;
An orientation sensor that detects the current orientation of the nose;
A yaw angle command is generated to control the current nose direction to the horizontal target direction by changing the rotation speed of the upper main rotor and the lower main rotor, and the rotation speed of the pitching propeller is adjusted to the target point. A rotor control unit that generates a horizontal position command for controlling the horizontal position of the fuselage by controlling the horizontal thrust by the upper main rotor and the lower main rotor by changing in proportion to the horizontal distance,
Furthermore, the rotor control unit tries to turn in either one of the left and right directions by the cyclic pitch input of the upper main rotor that attempts to return the forward leaning posture to the horizontal posture when the aircraft is caused to fly forward by the pitching propeller. Control of the rolling moment by the cyclic pitch input to the blades of the lower main rotor, and the yawing moment of the aircraft by generating the torque in the yaw direction by changing the rotational speed of the upper main rotor and the lower main rotor A counter-rotating rotary wing aircraft characterized in that, by performing control, a turning force for turning the airframe in a direction opposite to the one direction is applied, and the left and right turning forces are balanced to perform straight flight. A current altitude detector that detects the altitude component of the spatial coordinates of the current location from the altitude signal from the altimeter attached to the aircraft,
A vertical displacement calculator that calculates an altitude difference based on the altitude component of the spatial coordinates of the target location and the altitude component of the spatial coordinates of the current location;
In addition to generating the yaw angle command and horizontal position command, the rotor control unit changes the rotational speed of the upper main rotor and lower main rotor to change the altitude command to control the aircraft's current altitude to match the target altitude. The contra-rotating rotorcraft according to claim 3, wherein the counter rotating rotorcraft is generated.

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