JPS6028002B2 - Steering control system for flying bodies - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION [Field of Industrial Application] The present invention relates to a control method for controlling the steering of a flying object such as a missile.

[Technical background of the invention; prior art]

In controlling a flying object, control by the control surface is usually
As shown in the control block diagram of FIG. 4, the control is performed by feeding back the aircraft state variable Q according to the control purpose.

For example, if you want to control the altitude of the flying object, the altitude is selected as the aircraft state variable q to be fed back, and if you want to control the attitude angle of the flying object,
The attitude angle is selected as the aircraft state variable Q to be fed back, and when it is desired to roughly control the acceleration of the flying object, acceleration is selected as the aircraft state variable Q to be fed back. The control system of many flying objects is the fifth
As shown in the control block diagram in the figure, attitude angular velocity feedback is provided for the purpose of improving attenuation of the aircraft attitude motion relative to steering.

Furthermore, in order to reduce the cost of the actuators provided in the control system, some use relay-driven actuators or adopt a relay steering control system, as shown in FIG. By the way, in the control performed by feedback of the aircraft state variables mentioned above, no matter what kind of aircraft state variables are controlled, the speed, altitude, weight, center of gravity, moment of inertia, and aerodynamic stability of the flying object are determined during flight. It is necessary to solve the problem that the response characteristics of the aircraft to steering also change because the

For example, the moment M6 (see FIG. 7) around the center of gravity of the aircraft, which is generated due to a steering angle of 6 in steering control, is expressed by the following equation.

M6=back pWS&Cm6×6/1 '11Here, M
6 = Moment around the center of gravity of the aircraft caused by taking the rudder angle 6 = Air density V = Flying body speed SB = Flying body reference area c = Flying body reference length Cm6 = Rudder effectiveness aerodynamic coefficient 6 = Rudder Angle 1 = Moment of inertia of the flying body.

Of these, except for the reference area SB and reference length c of the flying object itself, all other variables change during flight. For example, the air density p decreases as the altitude increases, and the flying body moment of inertia 1 also changes with fuel consumption. The speed V also normally increases or decreases due to acceleration and deceleration after the flying object is launched. The aerodynamic coefficient Cm6 also changes depending on the Mach number, that is, the speed of the flying body. Therefore, even if the steering angle is the same as 6,
As is clear from the above equation '11, for example, as the speed V increases, M6 increases.

That is, the moment M6 around the center of gravity of the aircraft generated by steering changes depending on changes in the speed, altitude, weight, center of gravity, moment of inertia, and aerodynamic stability of the flying object. Therefore, in a linear control system or a conventional relay control system, if this change in the aircraft steering response characteristics becomes a problem, this is detected by some method to avoid changing the characteristics of the control system as a whole.

For example, a dynamic pressure sensor detects a dynamic pressure of 1/2oV2, and if the dynamic pressure becomes high, the static gain of the actuator is lowered to reduce the steering angle 6 by 4 degrees and keep M6 constant. It has been adopted. [Problems to be Solved by the Invention] In such a conventional method, when a change in the aircraft steering response characteristic becomes a problem in steering control of a flying object, as described above, in order to detect the aircraft steering response characteristic, This requires a special sensor and a special gain change mechanism to change the static gain of the actuator.

In general, a system in which the response characteristics of the entire control system change when the characteristics of the controlled object (in the above case, the aircraft steering response characteristics) change is said to be "without an adaptive control system." Conversely, when the characteristics of the controlled object change, the response characteristics of the entire control system change. A system in which the responsiveness of the entire control system does not change is called an ``adaptive control system'' or an ``adaptive control system.''

Detecting changes in the characteristics of a controlled object, such as detecting dynamic pressure, is called "identification," and adjusting the static gain of a controller, such as an actuator, based on the identification result is called "adaptive control operation." ” or “adaptive adjustment behavior.” In order to provide adaptive controllability, many conventional adaptive control systems use special sensors (identifiers) for detecting aircraft steering response characteristics, as shown in the control block diagram of FIG.
A special actuator static gain change mechanism (gain changer) is provided. This means that the hardware of the control system will be complex, and if it requires high precision, high reliability and maintainability will be required, and the cost of the hardware will also be high. There were some problems, such as not being able to get it.

[Means for solving problems]

The present invention was proposed to address the above-mentioned problems.In order to obtain the above-mentioned adaptive controllability, the present invention uses a special sensor (identification device) and a special actuator static gain change mechanism (adaptive control operation) as in the past. It is an object of the present invention to provide a steering control method for an air vehicle that is resistant to changes in aircraft steering response characteristics and has a certain degree of adaptive controllability without the need for a device.

To this end, the present invention utilizes a relay control system with normal attitude angular velocity feedback, and provides a phase lead compensation circuit in the attitude angular velocity feedback loop to generate a predetermined steady-state eye excitation vibration. yield within the loop,
The amplitude of this self-excited vibration is used to detect changes in aircraft steering response characteristics.
The invention is also characterized in that the flying object is steered by an adaptive control operation generated in a relay circuit or a similar nonlinear circuit by steady eye-excited vibrations. [Example] Hereinafter, an example of the present invention will be specifically described with reference to the drawings.

FIG. 1 shows a control block diagram of an embodiment according to the present invention, and this control system is composed of a Leray circuit or a Leray circuit that produces a positive or negative output when a given signal exceeds a required value. A nonlinear circuit 2 with an equivalent function, an attitude angular velocity 8 generated in the aircraft motion 4 at a steering angle 6 due to the operation of the control surface drive actuator 3 and the actuator 3 driven by the output of this nonlinear circuit 2, and the aircraft state variable. It has sensors 5 and 6 that respectively detect Q and output feedback signals.

In other words, the control system according to the present invention normally uses a feedback loop of the airframe state variable Q, as is found in many relay control systems for airborne vehicles, and further improves the attenuation of the attitude angle movement relative to the steering. It has a feedback loop with an angular velocity of 8, and in this respect is the same as that of the subordinate control system, but in the present invention, in particular, during the feedback loop of the attitude angular velocity P, For the purpose of inducing eye-excited vibration, a phase lead compensation circuit 7 is provided.
is provided, and the signal of the steady eye excitation vibration given by this phase lead compensation circuit 7 and the aircraft state variable Q which is negatively fed back are
The feedback signal and the input command signal Qc are applied to the nonlinear circuit 2 via the adder circuit 1.

By the way, the steady eye excitation vibration generated in the phase lead compensation circuit 7 differs from the steady eye excitation vibration of a normal relay control system in the following points.

The steady eye excitation vibration of a normal relay control system is, for example, in the case of a flying object, since the natural frequency of the airframe is fixed, the steady eye excitation vibration is also an eye excitation frequency close to that natural frequency. becomes.

This tends to result in low-frequency, large-amplitude vibrations, and for this reason, in a normal relay system, the width of the relay on and off is different, and in the short-flying gymnastics rudder system, the rudder width is controlled by the on and off of the relay. For example, turn on the steering angle by 10 degrees,
This will limit the steering angle to 2 degrees on and 2 degrees smaller than off. Due to this limitation, if the steering amount is reduced to reduce the steady eye excitation vibration amplitude, if a disturbance such as a gust of wind occurs that cannot be suppressed by the reduced steering amount,
My body becomes uncontrollable when I try to fly. On the other hand, the frequency of the steady eye excitation vibration generated in the present invention is a frequency that is appropriately higher than the natural frequency of the aircraft body caused by the phase lead compensation circuit 7,
For example, since the sound is approximately 1, the amplitude of the steering angle and the amplitude of the angular velocity of the aircraft attitude are small, and the amplitude of the aircraft attitude angle is extremely small and can be said to be not moving from an engineering perspective.

This is because the constant eye excitation vibration is generated at a frequency that is sufficiently higher than the frequency component to which the aircraft can respond. This constant eye excitation vibration component just before being input to the relay is expressed as the ot of the mother in,
Assuming that $in st is the steering signal up to a frequency at which the aircraft sufficiently responds from an engineering point of view, that is, near the natural frequency of the aircraft, the input signal to the relay can be expressed as follows.

ASjn St 10BSin.

t ■However, the o of the mountain s《,
Let A<<B. Then, as shown in FIG. 2, the gain of the steering command signal rudder in of the relay relative to st changes depending on the magnitude of the steady eye excitation vibration amplitude B.

In other words, if the magnitude of the steady eye excitation vibration amplitude B increases,
The equivalent gain of the relay for st of the steering command signal 瓜in becomes smaller, and conversely, if the steady eye excitation vibration amplitude B becomes smaller, the equivalent gain of the relay for st of the steering command signal ship jn becomes larger. Therefore, the phase advance circuit 7 is provided in the feedback loop of the attitude angular velocity P to incorporate this steady eye excitation signal component, which has a high frequency and a small amplitude, in the attitude angular velocity P.
The amplitude of the steady eye excitation vibration is amplified using the gain characteristics and static gain of this circuit.

Now, suppose that while flying, the speed, altitude, weight, center of gravity, and moment of inertia of the flying object change, and the aircraft steering response characteristics change.The aircraft is vibrating at a steady eye excitation vibration frequency and has a constant rudder angle amplitude. The attitude angular velocity response amplitude also changes.

For example, if a characteristic change occurs in which the speed increases and the response of the aircraft to constant steering increases, at the same time, the attitude and angular velocity of the aircraft also increases the constant eye excitation vibration amplitude. In this case, the output steady eye excitation vibration amplitude of the phase lead compensation circuit 7 also increases, so that the equivalent gain of the relay for the low frequency steering command signal component decreases. In other words, if we focus on the low-frequency steering command signal component, if the aircraft steering response characteristic becomes large in response to the same steering angle, the steady eye excitation vibration amplitude will increase and the equivalent gain of the relay will decrease, resulting in a constant response. The characteristics will be maintained. Expressed using the concept of adaptive control theory, changes in the characteristics of the aircraft are detected using the constant vibration amplitude,
The static gain of the actuator is changed using the relay's equivalent gain change characteristic caused by the amplitude characteristic. In reality, the phase lead compensation circuit 7 is designed to prevent noise signals with a frequency far higher than the frequency of the steady eye excitation vibration from being mixed into the output of the attitude angular velocity sensor 5, as well as the capacitors and resistances that form the basis of the circuit. In order to improve the selectivity of
When expressed as a transfer function, the denominator is quartic and the numerator is quadratic, as shown in the equation below.

・ ka (Saichi s20ru issue S+1) X total) (excellent) ‘3’ Here, ka=constant n = compensation circuit natural frequency = Compensation circuit damping constant 7 = phase lead compensation circuit constant 72= Same as above Ding 3= Same as above 74= Same as above S = Laplace transform parameter It is.

FIG. 3 is an example of a step command input response when the present invention is applied to a certain unmanned guided flight vehicle.

Here, the controlled object state variable is the vertical or horizontal acceleration of the aircraft. As can be understood from this figure, steady eye-excited vibration with a frequency of about 11 Hz and a total amplitude of about 0.3 seconds appears in the aircraft acceleration response, but the response at the center of the vibration can be regarded as the response of a first-order lag system. , and the time constant in this case is approximately 0
．． It is 4 seconds. In this example, no steering angle feedback is provided so that the center of the vibrating steering angle can take any desired steering angle within the movable steering angle range. [Effects of the Invention] As explained above, the flying object steering control system according to the present invention is basically a relay control system with normal attitude angular velocity feedback, so even a simple servo system is sufficient. In addition, a phase lead compensation circuit is installed in the attitude angular velocity feedback loop to induce the required steady eye excitation vibration in the loop, and the amplitude of this eye excitation vibration changes the aircraft steering response characteristics. , and the flying object is controlled by the adaptive control action generated in a relay circuit or similar nonlinear circuit by steady eye-excited vibration.In order to obtain the above-mentioned adaptive controllability, There is no need to provide a special sensor (identification device) such as a button or a special actuator static gain change mechanism (adaptive control operation device), so according to the present invention, the hardware of the control system can be simplified. Not only can the cost of hardware be reduced, but also advantages such as excellent reliability and maintainability can be obtained.

[Brief explanation of drawings]

1 to 3 are explanatory diagrams of embodiments of the present invention,
FIG. 1 shows a control block of an embodiment of the present invention. Fig. 2 is a characteristic diagram of the equivalent gain change of the relay caused by the steady eye excitation vibration amplitude, Fig. 3 is a diagram showing the step command input response of an application example of the present invention, and Figs. FIG. 4 is a block diagram of a feedback control system using a normal control surface; FIG.
Figure 5 is a block diagram of a system in which attitude angular velocity feedback is provided in the control system in Figure 4, Figure 6 is a block diagram of a system in which a relay is used for the control surface drive method in the control system in Figure 5, and Figure 7 is a block diagram of a system in which a relay is used for the control surface drive method in the control system in Figure 5. The figure shows the moment M around the center of gravity of the aircraft caused by the rudder angle.
6 and FIG. 8 are block diagrams when the control system of FIG. 5 and the adaptive control system are used. In Figure 1, 1...
-Addition circuit, 2...Nonlinear circuit, 3...Actuator, 4...Aircraft motion, 5...Sensor, 6
...Sensor, 7...Phase lead compensation circuit. pine convict pine 4th 8 legs 2 explanations bu 5 maps travel map 7 maps 8 solutions

Claims (1)

[Claims] 1 In a device that controls the steering of a flying object, A. A relay circuit that produces an output in the positive or negative direction when a given signal exceeds a required value, or a nonlinear circuit with an equivalent function, and B. A control surface drive actuator driven by the output, C a sensor that detects the attitude angular velocity and airframe state variable generated by the operation of the actuator and outputs a feedback signal, and D a phase lead provided in the attitude angular velocity feedback loop. A compensation circuit; B. An addition circuit that adds the input command signal, the negative feedback signal, and the steady self-excited vibration signal provided by the phase advance compensation circuit to the nonlinear circuit. A steering control system for a flying body featuring: