JPH06329093A - Control device - Google Patents

Abstract

(57) [Summary] [Purpose] An object of the present invention is to obtain a control device capable of stably controlling the aircraft as instructed even if the dynamic characteristics of the aircraft greatly change due to flight conditions. [Configuration] An inertial device 4 for detecting the motion of the airframe, and based on the information, an attitude angle with respect to the mainstream linear velocity is calculated by an attack angle calculation unit 5 and a roll attitude angle calculation unit 8, and an aerodynamic table unit 6 The compensation gain calculated by the control compensator gain calculation unit 7 from the data is set in the control compensator 1. [Effect] Since the attitude angle with respect to the mainstream linear velocity is calculated, and the aerodynamic derivative is used to set the control compensator gain, there are few restrictions on the attitude angle with respect to the mainstream linear velocity.
It is possible to realize a control device with little deterioration in stability and performance even when the flight conditions change, and it is possible to stably control the aircraft as instructed.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a control compensator used for stably operating a flying vehicle or an aircraft body as instructed.

[0002]

2. Description of the Related Art FIG. 9 is a block diagram of a control device using a conventional control compensation gain setting method. In FIG. 9, 1 is a control compensator, 2 is a steering device for moving a steering wing,
Reference numeral 3 is a body motion characteristic which is a control target of the present invention, and reference numeral 9 is an inertial sensor for detecting acceleration and angular velocity of the body.

The conventional control device controls the speed at which the aircraft is launched from the mother aircraft in order to cope with changes in the dynamic characteristics and aerodynamic characteristics of the airframe due to the flight of the flying body and the aircraft at a wide range of speeds and altitudes. By setting the control compensation gain according to the altitude, the response to the command was kept constant. For example, when the flying body is flying at a constant speed, the higher the altitude, the smaller the aerodynamic force acting on the aircraft and the slower the response of the aircraft. Therefore, when the altitude becomes higher, the control compensation gain is increased to increase the control amount and make the response to the command constant.

[0004]

In the control device using the conventional control compensator as described above, compensation is made only for the dynamic characteristics and aerodynamic characteristics of the airframe that change depending on the speed and altitude of the airframe. It does not take into consideration that the aircraft dynamic characteristics change depending on the attitude angle defined by the angle between the main streamline velocity and the aircraft axis. The body dynamic characteristics change depending on the posture angle, for example, the aerodynamic force increases as the posture angle increases, or the turbulence of the nearby air is generated due to the influence of the body. Therefore, for example, in the case of a large angle of attack, there is a problem that the body motion cannot be compensated.

The present invention has been made to solve the above problems, and not only the speed and altitude of the aircraft,
By setting the compensation gain of the control compensator according to the angle of attack and the roll attitude angle, the object is to make it possible to stably compensate the aircraft motion even if the attitude angle with respect to the main stream velocity of the aircraft changes.

[0006]

In the control device according to the present invention, first, the velocity information from the inertial device 4 is used to calculate the aircraft attitude angle with respect to the main streamline velocity, and the wind tunnel test is performed in advance to measure the aircraft attitude angle. Select an aerodynamic derivative coefficient according to the attitude angle, and calculate the control compensator gain using the aerodynamic derivative coefficient according to the attitude angle so that stable aircraft motion control can be performed regardless of the attitude angle of the aircraft Then, the compensation gain is set momentarily.

[0007]

By using the control device as described above,
Even when there is a large attitude angle with respect to the main stream velocity, it is possible to control the aircraft stably according to the command.

[0008]

【Example】

Embodiment 1 Hereinafter, an embodiment of the present invention will be described with reference to FIG. Figure 1
FIG. 3 is a block diagram showing an embodiment of a control device of the present invention. In FIG. 1, 1 to 3 are the same as the above conventional device. However, 1 means that the set value changes from moment to moment according to the present invention. Reference numeral 4 is an inertial device that detects the body motion with an inertial sensor and calculates the position, altitude, speed, and attitude angle.
5 is an angle of attack calculator, 6 is an aerodynamic table unit, and 7 is a control compensator gain calculator.

In the block diagram of FIG. 1, the velocities (X-axis velocity U, Y-axis velocity V, Z-axis velocity W) and main streamline velocity V _M of the orthogonal body 3 axes calculated by the inertial device 4 are utilized. The angle-of-attack α of the aircraft is calculated by the angle-of-attack calculation unit 5 using “Equation 1”. Then, the aerodynamic differential coefficient corresponding to the aircraft attack angle α is selected from the aerodynamic table unit 6. The aerodynamic table section 6 is shown in FIG.
An example is shown. Since the aerodynamic table unit 6 has the configuration shown in FIG. 2, it is possible to select the aerodynamic differential coefficient corresponding to the angle of attack. The control compensator gain calculation unit 7 calculates the compensation gain by using the aerodynamic differential coefficient selected in the aerodynamic table unit 6 in accordance with the attack angle coefficient and the attack angle, which are calculated in advance, by the "equation 2". By setting the compensator gain of this invention according to the attitude angle of the aircraft during flight,
It is possible to realize a control device that is not restricted by the attitude angle with respect to the mainstream linear velocity of the airframe.

The velocity calculated by the inertial device 4 is shown in FIG.
As shown in (3), it is the combined velocity of the velocity in the x-, y-, and z-axis directions and the velocity in the 3-axis direction. In addition, the angle of attack of the airframe calculated by the angle-of-attack calculation unit 5 indicates the angle formed by the x-axis velocity U and the main stream velocity V _M , as shown in FIG.

[0011]

[Equation 1]

[0012]

[Equation 2]

Embodiment 2 In the block diagram of FIG. 5, the velocities (X-axis velocity U, Y-axis velocity V, Z-axis velocity W) and main streamline velocity V _M of the orthogonal three axes of the machine body calculated by the inertial device 4 are shown. Utilizing this, the roll attitude angle φ of the machine body is calculated by the mathematical expression 3 in the roll attitude angle calculation unit 8. Then, the aerodynamic differential coefficient corresponding to the roll attitude angle φ of the machine body is selected from the aerodynamic table unit 6. Aerodynamic table section 6
Has the same configuration as that of the first embodiment, and an aerodynamic derivative corresponding to the roll attitude angle can be selected. The control compensator gain calculation unit 7 uses the aerodynamic differential coefficient selected in the aerodynamic table unit 6 to calculate the compensation gain in accordance with the roll attitude angle coefficient and the roll attitude angle which are calculated in advance by "Equation 4". By setting the compensator gain of the present invention according to the roll attitude angle φ of the machine body during flight, it is possible to realize a control device that is not restricted by the roll attitude angle φ of the machine body.

The velocity calculated by the inertial device 4 is shown in FIG.
As shown in (3), it is the combined velocity of the velocity in the x-, y-, and z-axis directions and the velocity in the 3-axis direction. Further, the roll attitude angle calculated by the roll attitude angle calculation unit 5 indicates an angle with respect to the relative speed, as shown in FIG.

[0015]

[Equation 3]

[0016]

[Equation 4]

Embodiment 3 In the block diagram of FIG. 7, the velocities (X-axis velocity U, Y-axis velocity V, Z-axis velocity W) and main streamline velocity V _M of the orthogonal three body axes calculated by the inertial device 4 are shown. Use the aircraft's angle of attack α
Is calculated by the mathematical expression 5 of the attack angle, and the roll attitude angle φ is calculated by the mathematical expression 3 of the roll attitude angle. Then, the aerodynamic differential coefficient corresponding to the aircraft attack angle α and the roll attitude angle φ is selected from the aerodynamic table unit 6. Aerodynamic table section 6
Since has the same configuration as that of the first embodiment, it is possible to select an aerodynamic derivative corresponding to the angle of attack and the roll attitude angle. The control compensator gain calculation unit 7 uses the aerodynamic differential coefficient selected in the aerodynamic table unit 6 to calculate the compensation gain based on the coefficient and the attitude angle which are calculated in advance. By setting the compensator gain of the present invention according to the aircraft's attack angle α and roll attitude angle φ during flight, a controller that is not limited to the aircraft's attack angle α and roll attitude angle φ is realized. it can.

Embodiment 4 In the block diagram of FIG. 8, the velocities (X-axis velocity U, Y-axis velocity V, Z-axis velocity W) and main streamline velocity V _M of the orthogonal three machine axes calculated by the inertial device 4 are shown. Use the aircraft's angle of attack α
Is calculated by the mathematical expression 5 of the attack angle, and the roll attitude angle φ is calculated by the mathematical expression 3 of the roll attitude angle. Then, an aerodynamic differential coefficient corresponding to the aircraft attack angle α and the aerodynamic roll attitude angle φ is selected from the aerodynamic table unit 6. Since the aerodynamic table unit 6 has the same configuration as that of the first embodiment, the aerodynamic differential coefficient corresponding to the angle of attack can be selected. Further, the aerodynamic differential coefficient corresponding to the altitude and speed output from the inertial device 4 is selected from the aerodynamic table unit 6 having the same configuration as that of the first embodiment. Then, the control compensator gain calculation unit 7
At altitude, speed, aircraft angle of attack α and roll attitude angle φ, aerodynamic differential coefficient and angle of attack coefficient and angle of attack calculated in advance,
Calculated from roll attitude angle, altitude, and speed. By setting the compensator gain of the present invention according to altitude, speed, aircraft attack angle α and roll attitude angle φ, there are less restrictions than conventional control devices and stable even when flight conditions change. It is possible to realize a control device with less deterioration in performance and performance.

[0019]

As described above, according to the present invention, the attitude angle with respect to the main streamline velocity is calculated, and the control compensator gain is set by using the aerodynamic derivative corresponding to the attitude angle. It is possible to realize a control device with less attitude angle restriction and less deterioration in stability and performance even when flight conditions change, and it is possible to control the aircraft as instructed.

[Brief description of drawings]

FIG. 1 is a block diagram showing a first embodiment of the present invention.

FIG. 2 is a diagram showing an example of an aerodynamic table section of the present invention.

FIG. 3 is a diagram showing a coordinate definition of velocity according to the present invention.

FIG. 4 is a diagram showing a definition of an attack angle according to the present invention.

FIG. 5 is a block diagram showing a second embodiment of the present invention.

FIG. 6 is a diagram showing a definition of a roll posture angle according to the present invention.

FIG. 7 is a block diagram showing a third embodiment of the present invention.

FIG. 8 is a block diagram showing a fourth embodiment of the present invention.

FIG. 9 is a block diagram showing an example of a control device according to a conventional technique.

[Explanation of symbols]

 1 Control Compensator 2 Steering Device 3 Body Dynamics 4 Inertia Device 5 Angle of Attack Calculation Unit 6 Aerodynamic Table Unit 7 Control Compensator Gain Calculation Unit 8 Roll Attitude Angle Calculation Unit 9 Inertial Sensor

Claims (4)

[Claims] 1. An inertial device for detecting the acceleration and angular velocity of the airframe to calculate the position, velocity and attitude angle of the airframe, a control compensator for controlling the attitude angle of the airframe, and a steering device for moving a steering wing. Input the output of the inertial device, the angle of attack calculation unit that calculates the angle of attack from the mainstream linear velocity and the aircraft axis velocity, and the aerodynamic table that has the aerodynamic derivative coefficient inside and outputs the aerodynamic derivative coefficient according to the above attack angle. Section, a compensator gain is calculated using the aerodynamic differential coefficient corresponding to the angle of attack output in the aerodynamic table section and the angle of attack, and a control compensator gain calculation for setting a gain according to the angle of attack is calculated. And a control unit. 2. An inertial device that detects the acceleration and angular velocity of the airframe to calculate the position, velocity, and attitude angle of the airframe, a control compensator that controls the attitude angle of the airframe, and a steering device that moves the steering wings. The output of the inertial device is input, the roll attitude angle calculation unit that calculates the roll attitude angle from the main streamline velocity and the airframe axis speed, and the aerodynamic derivative coefficient is internally provided, and the aerodynamic derivative coefficient according to the roll attitude angle is output. To calculate the compensator gain using the aerodynamic table section, the aerodynamic differential coefficient corresponding to the roll attitude angle output by the aerodynamic table section, and the roll attitude angle, and set the gain according to the roll attitude angle. And a control compensator gain calculation section. 3. An inertial device that detects the acceleration and angular velocity of the airframe to calculate the position, velocity, and attitude angle of the airframe, a control compensator that controls the attitude angle of the airframe, and a steering device that moves the steering wings. Input the output of the inertial device and calculate the angle of attack from the mainstream linear velocity and the airframe axial velocity, and input the output of the inertial device to calculate the roll attitude angle from the mainstream linear velocity and the aircraft axial velocity. A roll attitude angle calculation section, an aerodynamic table section which has an aerodynamic derivative coefficient therein, and outputs an aerodynamic derivative coefficient corresponding to the attack angle and the roll attitude angle, and an attack angle and roll attitude output by the aerodynamic table section And a control compensator gain calculation unit for calculating a compensator gain using the aerodynamic derivative corresponding to the angle, the attack angle and the roll attitude angle and w, and setting the gain according to the aircraft attack angle and the roll attitude angle. It is equipped with Control device. 4. An inertial device that detects the acceleration and angular velocity of the airframe to calculate the position, velocity, and attitude angle of the airframe, a control compensator that controls the attitude angle of the airframe, and a steering device that moves the steering wings. Input the output of the inertial device and calculate the angle of attack from the mainstream linear velocity and the airframe axial velocity, and input the output of the inertial device to calculate the roll attitude angle from the mainstream linear velocity and the aircraft axial velocity. A roll attitude angle calculation unit that has an aerodynamic derivative coefficient inside, and an aerodynamic table section that outputs an aerodynamic derivative coefficient according to the attack angle and roll attitude angle and an aerodynamic derivative coefficient according to the speed and altitude of the airframe, and The compensator gain is calculated by using the aerodynamic differential coefficient corresponding to the angle of attack and the roll attitude angle and the aerodynamic differential coefficient corresponding to the speed and altitude output from the aerodynamic table section, the attack angle and the roll attitude angle, and the speed and altitude of the airframe. Calculate the speed of the aircraft And a control compensator gain calculation unit for setting a gain according to altitude, angle of attack, and roll attitude angle.