JPH02230097A - Guided missile - Google Patents

Abstract

(57) [Summary] This bulletin contains application data before electronic filing, so abstract data is not recorded.

Description

[Detailed Description of the Invention] [Industrial Application Field] This invention relates to the improvement of a guided flying vehicle, and more specifically, the invention relates to the improvement of a guided flying vehicle, and more specifically, the front wing and the main wing are moved forward after combustion of a rocket motor, and one control device is moved forward. It is characterized by the ability to improve the performance of the guided flying object by simultaneously steering the control vanes located in front and behind the main wing by sending control signals to two actuators each for the front and rear wings. be.

[Conventional technology]

Figures 6 and 7 are schematic diagrams showing examples of conventional guided flying vehicles. In the figures, (1) is the fuselage, (21
Before! i31 is the main wing, and (4) is the rear wing. Here, FIG. 6 shows an example of a fuselage using the front red steering system, and FIG. 7 shows an example of a fuselage using the rear 'R' steering system. Figure 8 is a schematic diagram showing the internal structure of a conventional guided flying vehicle (101 is the rocket motor, Xm1 is the center of gravity position of the aircraft before the rocket motor burns,
Xm2 is the center of gravity position of the aircraft after rocket motor combustion #
apl is the point of force on the aircraft.

Figure 9 shows an example of the state of the front wing steering system at the start of a turn.
Figure 10 is a schematic diagram showing an example of the state during steady turning using the same system. The orthogonal line (8b) is the y-axis.In the figure, the arrow δ indicates the amount of steering angle of the steering blade, the arrow α indicates the amount of angle of attack, and the white arrow indicates the direction of the mainstream.

In the aircraft shown in Figure 8, the center of gravity before the rocket motor burns is Xml, but as the rocket motor burns after the aircraft starts flying, the weight at the rear of the aircraft becomes lighter. The center of gravity moves forward to Xm2.

As a result, the distance from the center of gravity to the aircraft's point of impact is
It changes from Δx1 to Δχ2. That is, as the rocket motor burns, the center of gravity of the aircraft moves forward, the distance between the center of gravity and the point of impact becomes longer, and the aircraft becomes quieter and more stable.

Now, consider the case where turning acceleration is generated in the positive direction of the y-axis. At the start of a turn, a positive rudder angle is first applied to generate lift on the front wing, thereby generating a clockwise moment in FIG. 10 and rotating the aircraft in that direction. As the angle of attack increases, lift is generated from the main wings and fuselage, and the value of the clockwise moment also changes at the same time. In the case of a statically stable aircraft, the moment decreases as the angle of attack increases, so there is usually an angle of attack at which the moment becomes zero. This angle of attack is called the trim angle, and is usually uniquely determined for each aircraft once the rudder angle is determined. The state in which the angle of attack is maintained at the trim angle is called the trim state, and as shown in Figure 10, the aircraft is in the 1-rim state during a steady turn. At this time, the steering angle is set to a value corresponding to the trim angle required to generate the required turning acceleration.

Figures 11 and 12 are schematic diagrams showing an example of the state at the start of a turn and an example of the state during a steady turn, respectively, of the rear wing steering system, and the symbols in the figures are the same as in Figures 9 and 10. It is. However, the angle of attack α and the steering angle δ are positive if they are in the same direction as in FIG.

In order to generate a clockwise moment at the start of a turn, unlike the case of the front H steering system, it is necessary to use a negative steering angle to generate negative lift at the rear as shown in Figure 12. Become. In a statically stable aircraft, the moment decreases as the angle of attack increases, as in the case of the front N steering system, and during a steady turn, the aircraft enters a trim state as shown in Figure 12.

The steering angle at this time also takes a negative value, unlike in the case of the front wing steering method.

[Problem to be solved by the invention]

As shown in Figure 10, in the case of a canard steering system using a statically stable aircraft, the rudder angle during a steady turn is positive, so the effective angle of attack of the canard is the same as that of the aircraft. greater than the angle of attack. Therefore, the upper limit of the trim angle that can be taken is constrained by the stall characteristics of the front wing, and the maximum turning acceleration during steady turning is also constrained by this.

On the other hand, with the rear wing steering system, as shown in Figure 12, acceleration occurs in the opposite direction to the intended direction at the start of a turn, which deteriorates the aircraft's responsiveness.

In either steering method, as shown in Figure 8, after the start of flight, the center of gravity moves forward as the rocket motor burns, and the distance between the center of gravity and the point of impact of the aircraft becomes longer. This increases static stability and reduces the aircraft's responsiveness.

This invention was made to solve this problem. After the rocket motor burns, the front wing and main wing are moved forward, and two types of control wings are simultaneously steered to improve performance. It proposes the body.

[Means to solve the problem]

The means for improving the performance of a guided flying object according to the present invention is to move the front wing and the main wing forward by a front wing moving mechanism and a main wing moving r# mechanism, respectively, after combustion of the rocket motor.
By sending a control signal from one control device to each actuator of the steering vanes attached to the front and rear of the main wing, these actuators are actuated simultaneously.

[For production]

The guided flying object in this invention has two types of steering wings: a front wing in front of two main wings and a rear wing behind the main wings, and these wings receive control signals from one control device to control each wing. The vehicle is steered by simultaneously operating two actuators.

By moving the front wings and main wings forward, the point of force of the aircraft moved forward, making it possible to shorten the distance between the center of gravity and the point of force, which had become longer due to the combustion of the rocket motor.
Static stability is reduced and aircraft responsiveness is improved.

In addition, since two types of steering blades are simultaneously steered by one control device, the response of the front H steering system is faster and the rear
It would be possible to obtain the advantages of both methods, such as the ability to obtain a high trim angle of the steering system, and to improve the performance of the training aircraft.

〔Example〕

FIG. 1 is a schematic diagram showing an embodiment of the present invention. The dashed lines indicate the positions of the front wing and main wing before rocket motor combustion.

FIG. 2 is a sectional view showing the internal structure of the guided flying vehicle according to the present invention, and (5) is a control device for steering;
6a) is the actuator of the front wing (2) + (6b) is the actuator of the rear wing (4), and (7) is the control device {5
} control signal to the actuator (6b), (9a) is the front wing forward movement rilJ mechanism, (9b)
is the forward movement of the main wing! ! ll machine W, DO) is a rocket motor.

Figure 3 is a cross-sectional view showing the internal structure of the guided flying vehicle shown in Figure 1 during operation. This is the distance between the center of gravity and the point of impact after the wing and main wing have moved forward.The broken line indicates the position of the front wing and main wing before the rocket motor combustion.

4 and 5 are schematic diagrams showing an example of the operation of the guided flying vehicle shown in FIG. 1, and are respectively an example of the aircraft at the start of a turn after the front wing and main wing have moved forward, This is an example of the aircraft during steady turning. In the figure, (8) is the coordinate axis set within the plane of the figure, (8a) is the X axis, and (8b) is the y axis. Arrow δ indicates the steering angle of the steering blade, δa indicates the steering angle of the front wing, δb indicates the steering angle of the rear wing, and arrow α indicates the angle of attack.

In each figure, open arrows indicate the direction of airflow.

In a statically stable guided spacecraft configured as described above, the point of impact on the aircraft is at the rocket motor 0, as shown in Figure 2.
Since it is constant regardless of the combustion of 0), the rocket motor α■
With the combustion of , the distance between the center of gravity position and the point of force is
It changes in the direction of becoming longer from Δχ2. When this aircraft receives a signal indicating the completion of combustion from the rocket motor 00), the front H forward movement mechanism (9a) and the main wing forward movement mechanism (9b) operate as shown in Figure 3, and move to the position indicated by the broken line. Move the front wing and main wing to the position shown by the solid line in front.

By moving the front wings and main wings forward, the aircraft's landing point can be adjusted.
It moves forward to Xcp2, and the distance from the center of gravity Xml becomes shorter from Δx2 to Δx3. Due to the shortening of the distance between the center of gravity and the point of impact, the rocket 1 motor α
The static stability of the aircraft, which had increased with the combustion of ω, decreased,
This also has the effect of improving the responsiveness of the aircraft, which had been deteriorating.

At the start of a turn, as shown in Fig. 4, the front R (21) actuator (6a) operates in the direction of raising the leading edge of the front wing (2) according to the control signal from the control device (5). The actuator (6b) of the wing (4) operates in the direction of lowering the leading edge of the rear wing (4).At this time, the aircraft moves toward the front wing (2).
) is steered, the aircraft's responsiveness improves, and it is possible to prevent or reduce the occurrence of adverse acceleration in the rear wing steering system, which has the same effect as steering only the front wings.

During a steady turn, as shown in Fig. 5, the actuator (6a) of the front N (2) operates in the direction of raising the leading edge of the front 1g (2) according to the control signal from the control device (5). The actuator (6b) of the rear ``I (4)
4) It operates in the direction of lowering the leading edge. At this time, the clockwise moment required to balance the aircraft is
And after 1i! 41, the amount of steering angle for each wing can be smaller than when only the IM type steering wing is steered.

By requiring a small positive steering angle of the front wing (2), restrictions on the angle of attack caused by the stall characteristics of the H are relaxed.
High trim angles are possible. In addition, by reducing the negative steering angle of the rear wing, the amount of acceleration generated in the opposite direction to the direction that the aircraft is originally trying to obtain can be kept to a minimum.
In addition, the advantage of the rear wing steering system is that a high trim angle can be achieved, which has the effect of allowing a larger maximum acceleration to be obtained.

〔Effect of the invention〕

As explained above, this invention improves the responsiveness of the aircraft by moving the front wing and main wing forward after the rocket motor burns, and then increases the Send control signals for two control devices,
By simultaneously steering both wings, it is possible to obtain the advantages of both the front VT steering system and the rear wing steering system.

[Brief explanation of the drawing]

FIG. 1 is a schematic diagram of a guided vehicle according to an embodiment of the present invention, FIG. 2 is a sectional view showing the internal structure of a guided vehicle according to the present invention, and FIG. 3 is a schematic diagram of a guided vehicle according to the present invention. 4 and 5 are schematic diagrams showing an example of the operation of the guided spacecraft according to the present invention, and FIGS. 6 and 7 are cross-sectional views showing the internal structure of the conventional guided spacecraft during operation. Figure 8 is a schematic diagram showing part of the internal structure of a conventional guided flying vehicle, and Figures 9, 10, 11, and 12 are diagrams showing the operation of a conventional guided flying vehicle. It is a schematic diagram showing an example. In the figure, (2) is the front wing, (3) is the main wing, (4) is the rear wing, (5) is the control device, and (6) is the actuator 2 (9).
) is a moving mechanism. In each figure, the same reference numerals indicate the same or equivalent parts.

Claims (1)

[Claims] It has a guidance part for guiding the aircraft to the target and a control part for controlling the attitude of the aircraft, and the guidance part that flies by obtaining thrust from a rocket motor has a main wing and a control part installed in front of the main wing. a front wing, a rear wing provided behind the main wing, a control device for steering, a movement mechanism for moving the front wing position of the aircraft forward, and a movement mechanism for moving the main wing position of the aircraft forward. 1. A guided flying object, comprising: a movement mechanism for moving; and an actuator for the front wing and an actuator for the rear wing that operate in response to a control signal from the control device.