JPH059806B2 - - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a flying object guidance control device for guiding a flying object to a relatively low-speed target object such as a ground-to-ground missile, and more particularly to a device for compensating for deterioration in guidance accuracy of a flying object due to gravitational acceleration. .

Conventionally, proportional navigation has been widely used as a guidance navigation method for the above-mentioned flying object, and the proportional navigation and bank-to-turn method will be explained below.

In order to simplify the explanation, a two-dimensional drawing will be used. In FIG. 1, 1 is a flying object and 2 is a target object, and let m and t be the velocity vectors of the flying object 1 and the target object 2, respectively. Line of sight is the line connecting the flying object 1 and the target object 2.
Sight) 4. A reference line 3 fixed in an inertial coordinate system is taken, and the path angle of the flying object 1 and the angle of the target line 4 measured from this line are γ and σ, respectively. The flying object 1 uses air force, thrust, etc. to track the target object 2 with an acceleration vector _n perpendicular to the direction of travel. In proportional navigation, this acceleration vector _n
Control is performed so that the size of a _n is set as a _n =NeVcσ〓 (1). Here, Vc is the approach speed between the flying object 1 and the target object 2, and Ne is what is called an effective navigation coefficient. As a result, the path angle change rate γ of the flying object 1 is proportional to the visual angle change rate σ,
As long as we maintain this relationship, we will definitely fly Body 1 is the target body 2
We will meet at. Figure 2 shows this situation. If the flying object 1 and the target object 2 are at positions M and T, respectively, and the target object 2 moves at a constant speed on the TP, the flying object 1 has the MP Triangle PMT on top
They move straight while maintaining a similar relationship and meet at point P.
When the target body T ₁ starts an avoidance action and accelerates on P′T ₁ , if the flying body accelerates from M ₁ according to equation (1), the trajectory of the flying body is
They will travel on P′M ₁ and meet at point P′.

Next, the usual four-blade steering system and bank-to-turn system will be explained. FIG. 3 is an example of a guided flying object that performs wing steering. In the figure, the X-axis, Y-axis, and Z-axis are a coordinate system fixed to the aircraft body, with the X-axis in the direction of travel, and the Y-axis and Z-axis perpendicular to this in the right-handed system. The fuselage has two pairs of steering wings 9 and 10, which are mounted symmetrically with respect to the X-axis in the X-Y plane and the X-Z plane, respectively. 1 in the diagram
1 is a target detection device that detects the direction and position of a target using radio waves and infrared rays. In order to change the course of the aircraft based on these signals, the rudder blade 9 is tilted with respect to the X-Y plane, and the rudder blade 10 is tilted with respect to the X-Z plane.
Then, as shown in Figure 4, a lift force L is generated on the rudder blade in the airflow, and this L generates a moment around the center of gravity 12 of the aircraft as shown in Figure 3, and as a result, the entire aircraft including the wings It takes an angle α (called the angle of attack) with respect to the airflow, and lift is generated throughout the aircraft. The aircraft changes its course in the direction determined by this lift, so in the four-blade rudder system, the control blades 9, 1
By operating 0, you can guide the aircraft in any direction. Below, the X axis is the roll axis, and the Y
The axis is called the pitch axis, the Z axis is called the yaw axis, and
The rotation angles around the respective axes are respectively called roll angle, pitch angle, and yaw angle, and the steering blades 9 and 10 are called pitch rudder blade and yaw rudder blade, respectively. The roll angle is controlled, for example, by steering the left and right wings of the pitch to opposite sides, thereby generating a moment about the roll axis. As shown in FIG. 5, when the pitch-axis and yaw-axis direction components of the change rate vector σ of the eye sight angle σ measured from the temporary inertial reference axis 3 are σ〓y and σ〓z, respectively, To guide a flying object using proportional navigation, each of the acceleration vectors y
The pitch rudder blade and the yaw rudder blade may be controlled so that the components a _ny and a _nz in the axis and z-axis directions satisfy a _ny = NeVcσ〓y …(2) a _nz = NeVcσ〓σz …(3) respectively. . Figure 6 is a view of the flying body from the rear.

Since the bank-to-turn system has only one pair of rudder blades, if it were to be a pitch rudder blade, it would not be able to move in the y direction. So, for example, if you want to move to the upper left, as shown in Figure 6a, first set the left and right rudder blades of the pitch to positive and negative positions, respectively.
After rotating to the left, as shown in Fig. 6b, the left and right pitch rudder blades are held straight and the robot moves in the desired direction. When proportional navigation is used, the direction of the σ vector is made to coincide with the y-axis by rotating the aircraft, and the pitch rudder is controlled so that equation (2) is satisfied while σ z becomes 0. FIG. 7 shows an example of a known flying object guidance control device. In the figure, 13 is the approach speed obtained from the Doppler radar, 14 is the navigation gain, 15 and 16 are the compensation circuits, 17 is the G bias signal, 18 is the amplifier, 19 is the limiter, 20 is the steering servo, and 21 is the aircraft transmission. Represents a function. Such a guidance control device can execute calculations based on equations (2) and (3), and control the pitch rudder based on these calculations. In this case, the G bias signal 1 is applied to the output signal of the compensation circuit 15 of the guidance control device shown in FIG.
The control signal is corrected by adding 7. This is because if only the control signals from equations (2) and (3) are used, the flying object will sink due to gravitational acceleration and the guidance accuracy will deteriorate. In other words, as shown in Figure 8, in the bank-to-turn system, when the lifting surface (X-Y) plane of the aircraft is tilted with respect to gravity, the Y-direction component Wy causes acceleration in the Y-direction, resulting in sideslip. . Therefore, the influence of this gravitational acceleration is compensated for by the G bias signal. However, since the magnitude of the G bias signal to be applied changes depending on the attitude of the flying object with respect to the direction of gravity, it is necessary to detect the direction of gravitational acceleration relative to the flying object by some method. For this purpose, conventionally, command signals were sent from the aircraft or ground base from which the flying vehicle was launched.

Therefore, an object of the present invention is to calculate signals from a detecting means for detecting the roll angle φ of a gyro mounted on a flying object and a sideslip angle detecting means for detecting the sideslip angle β of the airframe, thereby detecting an external signal. It is an object of the present invention to provide a device that can output a G bias signal without any information. That is, the magnitude of the sideslip angle β becomes a function of the roll angle of the aircraft. The roll angle of the aircraft can be detected by a roll angle detection means such as a rate integrating gyroscope, but the error increases over time due to drift. Therefore, the G bias signal is obtained by correcting the roll angle based on the output of the sideslip angle detection means.The present invention will be explained in detail below using examples.

FIG. 9 is a circuit diagram showing an embodiment of the flying object guidance control device, particularly the G bias signal generator, according to the present invention. In the figure, 22 is a sideslip angle detector for detecting the sideslip angle β; As shown in Figure 10, it is attached to the flying object, and its output signal B
is supplied to one input side of a subtracter 25 of a correction circuit 24 via a noise cutting filter 23. Reference numeral 26 is a roll angle detector consisting of a rate integrating gyroscope, etc., which detects the roll angle φ _G , and is attached to the flying object, and its output signal is sent to one input side of the adder 27 of the correction circuit 24. Supplied. The adder 27 outputs a signal of the correction value φ ₀ obtained by adding the initial setting value and the output signal from the limiter 29 constituting the comparison control circuit 28 in an adder 30 and a signal of the roll angle φ _G. The signal of the estimated roll angle value φ ₁ (φ _G +φ ₀ ) output from the adder 27 is sent to the function signal generating circuit 3 consisting of a sine calculator 31 and an amplifier 32.
3 and then supplied to the other input side of the subtracter 25. The output signal β _E of the subtracter 25 corresponds to the difference between the output signal (Ksinφ ₁ ) from the function signal generation circuit 33 and the output signal from the filter 23, and
4. The comparison control circuit 28 includes, in addition to the limiter 29 and the integrator 34, a bang-bang element 35 with a dead band, a pulse generator 36, and an integrator 37.

38 is an output circuit, which is composed of a cosine calculator 39, an amplifier 40, and a pitch angle correction circuit 41.

Here, the sideslip angle β detected by the sideslip angle detector 22 indicates the angle that the oblique shadow of the wind vector relative to the flying object on the XY plane makes with the X axis, as shown in FIG. In FIG. 10, 42 is a guidance control device.

The operation of the flying object guidance and control device having the above configuration will be explained below. The output signal of the roll angle detector 26 is determined when the direction of gravity is known, such as when the flying object is attached to the mother aircraft or a ground base, or when the flying object becomes horizontal due to the dihedral effect after launch. Set at the point in time. At this point, the value of the correction value φ ₀ signal output from the adder 30 is set to 0 by initial setting. Also, from now on, the output of the roll angle detector 26 is set to the roll angle φ _G
The signal corresponds to Let φ ₁ = φ ₀ + φ _G , and the expected side slip angle β at roll angle φ ₁ is
Let _Ksinφ1 . In FIG. 9, after the output signal B〓 of the sideslip angle detector 22 is passed through a filter 23 to appropriately remove noise, the difference β _E between this signal and the output signal Ksinφ ₁ of the function signal generation circuit 33 is calculated by a subtracter 25. The calculation result is integrated by an integrator 34.

Now, assuming that the actual roll angle φ is larger than the estimated value φ ₁ , the above difference β _E is always positive, and when the integrated amount exceeds the dead zone of the bang bang element 35 with dead zone, one pulse is generated by the pulse generating circuit 36. Occur. This pulse generation simultaneously resets integrator 34 to zero. Note that when the integrator 34 is configured using a charging/discharging circuit consisting of a capacitor, the capacitor is discharged. On the other hand, after assigning a specific sign to this pulse, the integrator 37
Perform the integration with . By this operation, the signal value of the correction value φ ₀ increases, and the estimated value φ ₁ of the roll angle φ increases and approaches the true value. Correction can be made in the same way even when the roll angle φ is smaller than the estimated φ ₁ . Note that when the roll angle correction value φ ₀ exceeds ±90°, the limiter 27 is activated. Using the signal of the obtained roll angle estimate φ ₁ , the output circuit 38 calculates gcosφ ₁ (g is the magnitude of the gravitational acceleration and is set by the amplifier 40), and further adjusts the pitch angle as necessary. After performing pitch correction in the correction circuit 41, the signal is output as a G bias signal. In the above embodiment, a rate integrating gyro is used as the roll angle detector, but a gyro compass or other roll angle detection device may also be used. Also, the circuit from the integrator 34 to the limiter 29 in FIG. 9 is β _E
Any circuit that increases or decreases the value of the correction value φ ₀ with respect to the average value of φ 0 can achieve the same effect as the above embodiment.

Furthermore, instead of the integrator 34, it may be constructed of a counter that can be increased or decreased by input pulses, and a multiplier that multiplies the output of this counter by an appropriate gain.

As explained above, according to the present invention, the G bias signal is output by correcting the drift of the output from the roll angle detection means such as a rate integrating gyroscope provided on the flying object with the output from the slip angle detection means. As a result, the G bias signal can be easily obtained without any external information, and there is no need to send a command signal from the outside.

[Brief explanation of the drawing]

Figure 1 is a diagram of the relationship between a flying object and a target when tracking a target using proportional navigation, Figure 2 is an explanatory diagram of a meeting course using proportional navigation, and Figure 3 is a diagram of a guided flying object using wing steering. Explanatory diagram showing an example, 4th
The figure is an explanatory diagram showing the lift force generated on the steering blade. Figure 5 is an explanatory diagram showing the Y-axis and Z-axis components of the visual angle change rate vector σ. Figure 6 a and b are bank-to-turn. Fig. 7 is a block diagram showing an example of a typical guidance control circuit for a guided flying vehicle that uses the wing steering method. Fig. 8 shows the roll angle φ and the roll angle detected by the roll angle detector. φ _G and correction value φ ₀
FIG. ₉ is a circuit diagram showing an embodiment of the flying object guidance control device according to the present invention, and FIG. 10 is an explanatory diagram of the sideslip angle β. 1... Flying object, 2... Target object, 22... Skidding angle detector, 23... Filter, 24... Correction circuit, 25... Subtractor, 26... Roll angle detector, 27... Addition device, 28...comparison control circuit, 2
9... Limiter, 30... Adder, 33... Function signal generation circuit, 38... Output circuit.

Claims (1)

[Scope of Claims] 1. A flying object guidance control device including a bias signal generating means for outputting a bias signal for compensating for a decrease in flying object guidance accuracy due to gravitational acceleration, wherein the bias signal generating means comprises: a sideslip angle detection means for detecting a sideslip angle based on a wind speed vector relative to the flying object; a roll angle detection means for detecting a roll angle of the flying object; a function signal generation circuit that calculates an estimated value; and a correction value based on the difference between the side slip angle estimated value based on the output signal of the roll angle detection means and the output signal of the sideslip angle detection means, and the output of the roll angle detection means. 1. A flying object guidance control device comprising: a correction means for correcting a signal; and an output circuit for outputting the bias signal as a function of the roll angle signal corrected by the correction means. 2. The correction means sets the sum of the output signal of the roll angle detection means and the correction value as an estimated roll angle of the flying object, and calculates the sum of the output signal of the roll angle detection means and the correction value by the function signal generation circuit based on the estimated roll value. Claim 1, further comprising a comparison control circuit that controls the magnitude of the correction value in a direction that matches the sideslip angle estimated value and the output signal of the sideslip angle detecting means. Guidance control device for flying objects.