JP2863665B2 - Automatic flight equipment for rotary wing aircraft - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to the maintenance of the attitude of an airframe in a rotary wing aircraft, the maintenance / change of speed and altitude, the automatic turning,
The present invention relates to an automatic flight device that enables automatic control such as an automatic approach.

[0002]

2. Description of the Related Art Rotary wing aircraft are inherently unstable compared to fixed wing aircraft, and SAs are required to compensate for the instability.
S (Stability Augmentation)
2. Description of the Related Art Stability increasing devices having a damper function, such as a System (stability increasing system), are employed in various types of rotary wing aircraft. Some rotary wing aircraft also have an ASE (Automatic Stable) that maintains the attitude, speed, and altitude of the fuselage.
(equipment equipment: automatic stabilizer)
AF that can change the speed, altitude, altitude, and automatic approach
E (Automatic Flight Equipm
ent: automatic flight device).

However, the change of the flight state by the conventionally used AFE often takes a long time because the change is performed with a constant gain. To improve this, complicated methods such as gain scheduling are required.

[0004] An example of an automatic flight control device for a rotary wing aircraft is disclosed in Japanese Patent Application Laid-Open No. Sho 51-60397. As shown in this example, an automatic flight device generally includes a sensor that detects a current value of a flight condition such as an attitude, a speed and an altitude of an aircraft in flight, and a flight condition and a flight condition that a pilot intends to hold or change. Means for inputting the target value of the state, and calculating the target value and the current value of the flight state input by these means to control the four axes of pitch, roll, yaw, and collective pitch for the actuator of the rotary wing. Generate a signal.

However, the automatic flight control device for a rotary wing aircraft disclosed in this publication automatically and automatically moves to a predetermined hovering point from a control start point and to a predetermined hovering point. It aims at automatic flight control of a flight that stops without overshooting. This flight probably travels like a "float" between hovering points set at a certain distance vertically and horizontally on the vast ocean, hovering at a certain altitude above the sea surface at the hovering point, and receiving the hydrophone It appears that only the vessel was suspended so that it could enter the water, aiming to detect enemy submarines that could not be captured by radar.

Therefore, in this automatic flight control device, the remaining distance to the target hovering point is calculated, and control is performed so that the vehicle is decelerated from just before the hovering point and stopped accurately at the target point. The control start point (or previous hovering point) is calculated by integrating the speed detected by the speed sensor with an integrator to calculate the remaining distance.
From the current position to the current position. This integrator has an integrating function when the deviation between the set speed and the current speed is 5 knots or less, and functions as a simple amplifier when the deviation is 5 knots or more to generate an acceleration / deceleration signal.

Although the above-described control using a special integrator seems to be effective for the above-mentioned special purpose of moving between hovering points, the speed is rather increased in the general automatic flight of a rotary wing aircraft. When the deviation of the speed is large, it is desirable to perform a quick response, and when the deviation of the speed is small, it is desirable to perform control such that the control value does not overshoot and the stability is improved.

[0008] In addition, the current IF of a helicopter
R (Instrument Flight Rule:
In flight by the instrument flight rules) method, 50-60k
Flight below the minimum instrument flight speed of t (V _MINI , different for each model) is prohibited. The reason is V _MINI
This is because in the so-called backside region of the following speeds, the relationship between the speed and the required power is opposite to the speed region above V _MINI , and it becomes extremely difficult to steer, in combination with the inherent instability of the helicopter.

Therefore, when landing in the IFR system, V _MINI
Although the approach is performed at the above high speed, the descent angle needs to be small in order to keep the descent rate low.
Navigation support system ILS used for landing in FR
(Instrument Landing System
m), the approach at a shallow approach angle (about 3 °) and at a high speed (about 70 kt) is common, but in the case of approach at a low angle and at a high speed, sufficient space is sufficient. It is necessary to have a large facility such as ILS, so there is a runway and there is enough space, and landing at IFR can only be done at an airport with ground facilities. (See FIG. 7).

However, in this case, a high-speed fixed wing aircraft and a low-speed rotary wing aircraft are mixed above the airport, which results in traffic congestion.

[0011] Further, since the IFR operation system such as the ILS has been developed mainly for operation by fixed wing aircraft, it is not suitable for a rotary wing aircraft in terms of economy and efficiency.

Therefore, as an IFR operation method that does not hinder the operation of the fixed wing aircraft and makes use of the characteristics of the rotary wing aircraft, an air route for a low altitude rotary wing aircraft is set below the existing air route. Takeoff and landing can be done at a steeper angle than with fixed wing aircraft.

In this case, there is no portion overlapping the fixed wing aircraft in operation, and traffic congestion does not occur (a schematic diagram is shown in FIG. 8).

In the case of a heliport having no runway, there are many obstacles (buildings and the like) around the helipad and a shallow approach angle cannot be obtained in many cases. In this case, too, an IFR approach at a steep angle is required. In addition, a steep angle approach can reduce noisy areas on the ground.

As one of the systems capable of guiding such an approach at a steep angle, MLS (M
The development of the microwave landing system (ultra-high frequency landing system) has been advanced.

However, when the approach angle is made deeper, the descent rate increases as the approach angle is made deeper, unless the speed is reduced, which is very dangerous.

[0017] Therefore, it is necessary to perform the approach at a steep angle at a low speed at the same time, but this is a flight in the above-mentioned backside area, and the steering becomes very difficult.

Therefore, although navigation systems such as MLS have been developed so far,
There is no flight control device that controls the aircraft's motion and enables automatic landing on helipads at steep angles and at low speeds.

[0019]

SUMMARY OF THE INVENTION The present invention has been made in view of the above-mentioned problems of the conventional automatic flight system of a rotary wing aircraft, and has a purpose of maintaining the attitude of the body during flight, maintaining or changing the speed and altitude, automatic turning, and airport.・ Response of speed control is fast in an automatic flight device of a rotary wing aircraft that automatically controls four axes of pitch, roll, yaw and collective pitch according to flight conditions such as automatic approach to a heliport,
An object of the present invention is to provide an automatic flight device having excellent stability and an automatic flight device capable of performing an automatic approach at a low speed and a steep angle.

[Means for Solving the Problems]

According to the present invention, there is provided a sensor for detecting a current value of a flight state such as a posture of an airframe in flight, a rate of change of attitude, a speed, an acceleration, an altitude, and a rate of change of altitude. A means for inputting a flight condition and a target value such as a speed and an altitude to be changed; a means for calculating a flight state target value corresponding to the input flight condition and the target value; and a means for achieving the target value. Means for generating four-axis control signals of pitch, roll, yaw, and collective pitch from the current value of the flight state and the calculated target value. In an automatic flying device for a rotary wing aircraft that automatically controls the four axes according to flight conditions such as holding or changing, automatic turning, airport, automatic approach to a heliport, etc., As the attitude angle target value calculating means pitch attitude, the target value of the speed and the magnitude of the deviation between the present value,
If the deviation is large, a pitch attitude angle target value calculating means for speed change which increases acceleration / deceleration, and if the deviation is small, a pitch attitude angle target value calculation means for speed stabilization which reduces acceleration / deceleration. Angle target value calculation means,
The speed change pitch attitude angle target value calculation means derives a speed change pitch attitude angle target value from the speed obtained from the speed sensor and the speed target value calculated by the speed target value calculation means, The stabilizing pitch attitude angle target value calculating means is:
The problem can be solved by deriving a pitch stabilization target value for speed stabilization from the speed obtained from the speed sensor, the speed target value calculated by the speed target value calculation means, and the acceleration obtained from the acceleration sensor.

[0021]

According to the structure of the present invention, the response speed and the stability can be made compatible by greatly accelerating and decelerating when the speed deviation is large and decreasing the acceleration and deceleration when the deviation is small.

[0022]

DESCRIPTION OF THE PREFERRED EMBODIMENTS Embodiments of the present invention will be described below with reference to the drawings.

FIGS. 1 to 4 are block diagrams of a flight control system of a helicopter incorporating an automatic flight device according to the present invention. The figure shows a case in which a fly-by-wire that transmits an operation signal to the control system by electric wiring is used, and an example of a control system related to a pitch axis, a roll axis, a yaw axis, and a CP axis in order from FIG. Show.

Hereinafter, with reference to these drawings, holding of the posture, holding / changing of the speed / altitude, automatic turning,
A control method such as an automatic approach will be described.

When the pilot 1 is steering by itself, the switching units 101, 201, 301, 401 for switching the pilot steering signals δyp, δxp, δzp, δcpp and the control signals δy, δx, δz, δcp from the control unit are provided. As shown by a broken line in each figure, it is connected to the circuit of the pilot steering signal.

The pilot 1 inputs a target value through the control panel 2 of the automatic flight system, and when the control calculation is started, the connection is switched at the same time, and the steering is performed by a signal from the automatic flight system. .

This automatic flight system monitors the following numerical values as parameters for control by various sensors.

Θ: pitch attitude angle Φ: roll attitude angle ：: yaw attitude angle Q: pitch rate P: roll rate R: yaw rate U: speed (front-back direction) U ′: acceleration (front-back direction) V: Speed (lateral direction) V ': acceleration (lateral direction) H: altitude H': altitude change rate (elevation rate) And, from these parameters and target values, pitch axis (y), roll axis (x), yaw axis ( Z), a control operation is performed for each of the collective pitch axes (CP), and 4
Control signals (δy, δx, δz, δcp) are output to the respective axes.

The control method for each axis will be described below in detail. FIG. 1 shows a block diagram of the control of the pitch axis. In the control of the pitch axis, the speed U and the acceleration U ′ obtained from the speed sensor 102 and the acceleration sensor 103 and the speed target value Uref calculated by the speed target value calculator 104 according to the input from the control panel 2 are used. Pitch attitude angle (θ) target value calculators 105 and 106 derive a pitch attitude angle target value θref. In the present invention, the pitch attitude angle (θ) target value calculator is a speed change pitch attitude angle (θ) target value calculator 105 for quickly approaching the target value when the difference between the target speed and the current value is large. Speed stabilization pitch attitude angle (θ) for converging to the target value without overshooting when the difference between the target value and the current value is small
Two target value calculators 106 are provided. As a result, when the difference between the target speed and the current speed is large, the acceleration and deceleration are greatly increased, and when the difference is small, the acceleration and deceleration are reduced, whereby both the responsiveness and the stability can be achieved. Further, the outputs of the two target value calculators 105 and 106 are weighted and combined by the adder 107 according to the magnitude of the difference between the target speed and the current speed, so that the outputs of the two are not suddenly switched. The discontinuity does not appear.

When the attitude holding mode is selected or when the attitude needs to be held during the automatic approach, the pitch attitude target value calculator 108 directly calculates the target value θref of the pitch attitude angle regardless of the speed. And the switching unit 109
, The posture is maintained by switching the circuit.

From the calculated pitch attitude angle target value θref and the current pitch attitude angle θ obtained from the pitch attitude angle sensor 110, a deviation θerr of the pitch attitude angle is obtained. Rate Q
Is supplied to the pitch rate target value calculator 112.

In the pitch rate target value calculator 112, θ
Pitch target value Qr required from err and Q values
Calculate ef.

The target value Qref of the calculated pitch rate
The deviation Qerr of the pitch rate is obtained from the values of the pitch rate and the Q, and the value and the rate of change Q ′ of the pitch rate are supplied to the pitch axis steering amount calculator 113.

The pitch axis steering amount calculator 113 calculates the steering amount δy of the vertical cyclic pitch from the values of Qerr and Q ′.

The signal of δy is supplied to the actuator 114 in the pitch direction of the main rotor, whereby the body movement 115 around the pitch axis is performed to maintain or change the pitch attitude and speed.

FIG. 2 is a block diagram of control of the roll axis. Control panel 2 for roll axis control
The roll attitude angle target value calculator 203 calculates a roll attitude angle target value Φref for preventing sideslip from the conditions input by the above and the speed obtained by the speed sensor 202.

Further, as shown in a portion surrounded by a chain line in FIG. 2, a skid speed sensor 204 such as an air data sensor and a y-axis direction accelerometer 205 are provided, and the skid speed is set to 0 based on the values. In this manner, the roll attitude angle target value calculator 206 can also calculate the roll attitude angle target value.

Calculated Φref and roll attitude angle sensor 2
07, the deviation Φerr of the roll attitude angle is obtained from the current roll attitude angle Φ obtained from the current roll attitude angle Φ, and the value and the roll rate P obtained by the roll rate sensor 208 are supplied to the roll rate target value calculator 209. Is done.

In the roll rate target value calculator 209, Φ
The target value Pref of the roll rate required for eliminating err is calculated, and the deviation Perr of the roll rate is calculated from the current roll rate P, and the Perr and the rate of change P 'of the roll rate are calculated. It is supplied to the roll axis steering amount calculator 210.

From the values of Perr and P ', the roll axis steering amount calculator 210 calculates the steering amount δ of the lateral cyclic pitch.
Calculate x.

The signal of δx is supplied to the actuator 211 in the roll direction of the main rotor, whereby the body motion 212 around the roll axis is performed to maintain the roll posture /
Changes and turns are performed.

FIG. 3 is a control block diagram of the yaw axis. In the control of the yaw axis, the target value Ψref of the yaw attitude angle is calculated by the yaw attitude target value calculator 302 based on the conditions input on the control panel 2. At this time, as shown in FIG. 3, it is also possible to calculate the target value of the yaw attitude or the yaw rate in accordance with the fluctuation of the torque obtained by the torque sensor 303.

A deviation Ψerr of the yaw attitude angle is obtained from Ψref and the yaw attitude angle 得 obtained by the yaw attitude sensor 304, and the value of the deviation Ψerr and the current value of the yaw rate R obtained by the yaw rate sensor 305 are used as the yaw rate. -It is supplied to the rate target value calculator 306.

In the yaw rate target value calculator 306, Ψer
A target value Rref of the yaw rate required to eliminate Ψerr from r and R is calculated, and a deviation Rer of the yaw rate is obtained from the value and the current yaw rate R. It is supplied to the steering amount calculator 307.

In the yaw axis steering amount calculator 307, Rerr
And R ′, the steering amount δZ of the pedal is calculated, and the value is supplied to the actuator 308 of the tail rotor, whereby the body motion 309 around the yaw axis is controlled, and the azimuth is maintained.

FIG. 4 is a block diagram of control of the CP (collective pitch) axis. In the control of the CP axis, the altitude target value calculator 402 and the altitude change rate target value calculator 403 use the altitude target value Href and the altitude change rate target based on the altitude or the altitude change rate input from the control panel 2. The value H'ref is calculated.

The current altitude H and the altitude change rate sensor 405 based on the values (Href, H'ref) and the altitude sensor 404
The current altitude change rate H 'by the CP axis steering amount calculator 40
6

Here, as shown in FIG.
It is also possible to measure the acceleration in the Z-axis direction by the Z-axis direction acceleration sensor 407 in order to increase the damping, and use it as one of the control conditions.

The CP-axis steering amount calculator 406 calculates the C value required to achieve the target value from the various values supplied.
The operation amount δcp of the P lever is calculated. This signal is transmitted to the actuator 4 in both the pitch direction and the roll direction of the main rotor.
08, whereby the body motion 409 about the collective pitch axis is controlled, and the altitude is maintained or changed.

As described above, by controlling the four axes of the pitch axis, the roll axis, the yaw axis, and the collective pitch axis, the body attitude is maintained, and the speed / altitude is maintained / changed.
Automatic turning, automatic approach, etc. are performed.

Note that the block diagrams shown in FIGS. 1 to 4 are merely examples, and the same control can be performed using sensors other than those shown in the drawings.

In addition, MLS (Microwave Lan)
In the case of performing an automatic approach using a ding system, the control of the present invention can be realized by supplying an MLS signal to each target value calculator.

Next, the change of the target value in the automatic approach mode will be described. In a mode other than the automatic approach mode, the control target value does not change on the way, and the control is performed by one operation.

In the automatic approach mode, it is impossible to reach the final target by one operation, but the control is performed by means of changing the target value of the control sequentially from descent to hover near the ground. Is going.

Except for changing the target value, the automatic approach mode is basically the same as the other modes.

FIG. 5 shows an example of changing the target value in the automatic approach. This is an example of approach where the vehicle descends near the ground (about 100 ft) while maintaining a constant speed and descent rate, and decelerates by flare just before landing. It enters at a deeper angle (about 6 °) than the standard approach angle at IFR (about 3 °).

The target during the steady descent is to maintain the speed and the descent rate. At the same time as the start of the flare, the target is to change and maintain the pitch attitude during the flare and to prevent the body from rising. After the flare ends, the goal is to maintain a speed of 0 and a descent rate of 0, that is, to maintain stable hovering near the ground.

The roll axis and the yaw axis are aimed at preventing the sideslip and maintaining the azimuth during the execution of the automatic approach.

The change of the target value as described above is performed by each of the following target value calculators for calculating the target value based on the input from the control panel. Pitch axis: speed target value calculator 104, pitch attitude angle target value calculators 105, 106 Roll axis: roll attitude angle target value calculator 203 Yaw axis: yaw attitude angle target value calculator 302 CP axis ... Altitude target value calculator 402 and altitude change rate target value calculator 403 These mainly monitor the altitude H, and calculate and change final values such as speed and descent rate depending on the state. Is going. The timing of these changes reflects the pilot's operational know-how.

Next, the control calculation in the control computer will be described. In the present invention, fuzzy inference is used for control calculation. FIG. 6 shows the procedure of fuzzy inference.

The fuzzy inference is made up of membership functions 501 and 502 indicating the size of input / output and linguistic fuzzy rules 503 indicating the contents of control.

Then, the error (control deviation) between the flight state and the target value obtained by each sensor and the numerical value of the flight state are compared with the membership function 501 of the antecedent part of the rule, and the degree of conformity is calculated. The degree of conformity determines the rule to be applied and the degree of application, thereby determining the fuzzy output,
By defuzzifying 504 the value, a control signal is obtained.

When fuzzy inference is used for control, non-linear control can be easily realized by the shape of the membership function. In addition, since the control is made up of linguistic rules, it can be used to introduce pilot control know-how into control. Is easy.

It should be noted that in the switching section of the pilot steering signal and the signal from the control device shown in FIGS. 1 to 4, when the control device is performing the steering, the pilot inputs a value greater than a certain value such as a stick or a pedal. it by _{(δyp> δyp 1, δxp>} δxp 1, δzp> δzp 1 or δcpp
> Δcpp ₁ ) If the connection is switched (override), the control of the aircraft can be taken over from the automatic flight system at any time, so that sudden accidents such as near misses and delicate maneuvers at landing can be safely dealt with. can do.

[0065]

As described above, according to the automatic flight system of the present invention, it is possible to obtain a quick response and excellent stability of the body speed during the automatic flight.

[Brief description of the drawings]

FIG. 1 is a block diagram of control of a pitch axis of an embodiment of an automatic flight device of the present invention.

FIG. 2 is a block diagram of control of a roll axis according to the same embodiment.

FIG. 3 is a block diagram of control of a yaw axis according to the same embodiment.

FIG. 4 is a block diagram of control of a collective pitch axis according to the same embodiment.

FIG. 5 is an explanatory diagram showing an example of a target value change in an automatic approach.

FIG. 6 is an explanatory diagram illustrating a procedure of fuzzy inference.

FIG. 7 is an explanatory diagram showing the concept of a sharp angle low speed approach and a low angle high speed approach.

FIG. 8 is an explanatory diagram showing the concept of an operating method for separately describing upper and lower air routes of a fixed wing aircraft and a rotary wing aircraft.

[Explanation of symbols]

Reference Signs List 1 pilot 2 control panel 101, 201, 301, 401 switching unit 102 speed sensor 103 acceleration sensor 104 speed target value calculator 105 speed change pitch attitude angle target value calculator 106 speed stabilization pitch attitude angle target value calculator 108 pitch Attitude target value calculator 110 Pitch attitude angle sensor 111 Pitch rate sensor 112 Pitch rate target value calculator 113 Pitch axis steering amount calculator 114 Main rotor actuator (pitch direction) 115, 212, 309, 409 Body motion 202 Speed sensor 203 Roll Attitude angle target value calculator 204 Side slip velocity sensor 205 Lateral accelerometer 206 Roll attitude angle target value calculator 207 Roll attitude angle sensor 208 Roll / rate sensor 209 Roll / rate target value calculator 210 Roll Steering amount calculator 211 Main rotor actuator (roll direction) 302 Target value calculator for yaw attitude 303 Torque sensor 304 Yaw attitude angle sensor 305 Yaw rate sensor 306 Yaw rate target value calculator 307 Yaw axis steering amount calculator 308 Tail Rotor actuator 402 Altitude target value calculator 403 Altitude change rate target value calculator 404 Altitude sensor 405 Altitude change rate sensor 406 CP axis steering amount calculator 407 Z axis direction acceleration sensor 408 Main rotor actuator (collective) 501 Antecedent membership Function 502 Consequence Membership Function 503 Fuzzy Rule 504 Defuzzy

──────────────────────────────────────────────────続 き Continuation of the front page (56) References JP-A-56-85101 (JP, A) JP-A-52-77399 (JP, A) JP-A-57-178997 (JP, A) 45438 (JP, B2) (58) Field surveyed (Int. Cl. ⁶ , DB name) B64C 27/57

Claims (1)

(57) [Claims] 1. A sensor for detecting a current value of a flight state such as a posture of an airframe during flight, a rate of change of attitude, speed, acceleration, altitude, and a rate of change of altitude, and a flight condition to be held or changed by a pilot. Means for inputting target values such as speed, altitude, and the like; means for calculating target values of flight conditions corresponding to the input flight conditions and target values; and current values of the flight conditions to achieve the target values. Means for generating control signals for the four axes of pitch, roll, yaw, and collective pitch from the calculated target values, holding the attitude of the aircraft during flight, holding or changing the speed and altitude, automatic turning, An automatic flight device for a rotary wing aircraft that automatically controls the four axes according to flight conditions such as an automatic approach to an airport and a heliport. The calculation means includes a speed change pitch / posture angle target value calculation means for increasing the acceleration / deceleration when the deviation is large, depending on the magnitude of the deviation between the target value and the current value of the speed, and for acceleration / deceleration when the deviation is small. And a pitch stabilization target value calculating means for speed stabilization. The speed changing pitch / posture angle target value calculating means includes a speed obtained from a speed sensor and a speed target. A pitch attitude angle target value for speed change is derived from the speed target value calculated by the value calculation means, and the speed stabilization pitch attitude angle target value calculation means calculates the speed obtained from the speed sensor and the speed target value calculation. An automatic flight device that derives a target value of a pitch attitude angle for speed stabilization from a target speed value calculated by a means and acceleration obtained from an acceleration sensor.