JPH0244199A - Guided missile - Google Patents

Abstract

PURPOSE:To improve a performance of a guiding missile by a method wherein two types of operating wings, i.e., a front wing in front of a main wing and a rear wing before the main wing are provided, these operating wings may receive a control signal from one controller device and then actuators for each of the wings are operated simultaneously. CONSTITUTION:As a missle body receives a signal indicating a completion of combus tion of a rocket motor 10, a part of a rear end of the missle body indicated by a broken line is separated from the body. A force applying point of the body is moved forward to Xcp2 and a distance from a position of center of gravity Xm2 is made short from DELTAx2 to DELTAx3. A static stability of the body increased once as a combustion in the rocket motor 10 is performed is decreased and then a responding characteristic of the body is improved. IN case of turning of the body, an actuator 6a for a front wing 2 is operated to raise a front edge of the front wing 2 with a control signal of a controller device 5 and a rear actuator 6b for a rear wing 4 is operated to lower the front edge of the rear wing 4. At this time, the body shows a better responding characteristic due to an operation of the front wing 2 and thus an occurrence of an acceleration in an opposite direction in a rear wing steering system can be prevented or reduced.

Description

[Detailed Description of the Invention] [Field of Industrial Application] This invention relates to the improvement of guided flying vehicles, and more specifically, the rear end portions of two aircraft are separated from the aircraft after the rocket motors have been combusted, and one control unit is used. The runner-up feature is that the control signal of the device is sent to two actuators each for the front wing and the rear wing, thereby simultaneously steering the control vanes located forward and aft of the main wing, thereby improving the performance of the guided flying vehicle. It is something to do.

[Conventional technology]

6 and 7 are schematic diagrams showing examples of conventional guided flying vehicles. In FIG. 92, (l) is the fuselage, (2
) is the front wing, (3) is the main wing, and (4) is the rear wing. Here, FIG. 6 shows an example of a front wing steering type aircraft, and FIG. 7 shows an example of a rear wing steering type aircraft. FIG. 8 is a schematic diagram showing the internal structure of a conventional guided flying vehicle. Xcpl is the point of impact on the aircraft.

Figure 9 is an example of the state at the start of a turn using the front wing steering system.
Fig. 10 is a schematic diagram showing an example of the state during steady turning using the same method, and (8) is the next coordinate axis set within the plane of the figure, and the (8a,) t-X axis parallel to the main flow. , the (ab)ty axis is orthogonal to the X 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 combustion of the rocket motor is Xm1, but after the first aircraft starts flying, the weight of the rear part of the two aircraft becomes lighter as the rocket motor burns. The double center of gravity moves forward to Xz2.

The distance from the center of gravity to the point of impact of the aircraft is
It changes from Δx1 to Δx2. That is, as the rocket motor burns, the center of gravity of the aircraft moves forward, and the distance between the center of gravity and the point of impact becomes longer, increasing the static stability of the aircraft.

Now, consider the case where turning acceleration in the positive direction of the y-axis is generated. 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 two aircraft in that direction. As the angle of attack increases, lift is also generated from the main wing and fuselage, and the value of the initial clockwise moment also changes. For 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 next state that keeps the angle of attack at the trim angle is called the trim state. As shown in FIG. 10, the aircraft is in a trim state during a steady turn.

At this time, the steering angle is

It is set to a value corresponding to the trim angle required to generate the required turning acceleration.

FIG. 11 and FIG. 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 front are as follows.

This is the same as factor 9 and Figure 10. However, the angle of attack α and the steering angle δ are positive if they are in the same direction as in FIG.

K to generate a clockwise moment at the start of a turn.
This is different from the case of the front wing steering method, and as shown in FIG. 12, it is necessary to take a negative steering angle and generate a negative lift force on the rear wing. In a statically stable aircraft, the moment decreases as the angle of attack increases, as in the case of the front wing steering system, and during a steady turn, the aircraft enters a trim state as shown in FIG. 12. The steering angle at this time also has a different value from 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 steering angle during a steady turn is positive at 1 cFi, so the effective angle of attack of the canard is The horns also get bigger. Therefore, the upper limit of the value that can be taken as the trim angle is limited by the stall characteristics of the front wing, and the maximum turning acceleration during steady turning is also restricted accordingly.

On the other hand, in the rear wing steering system, as shown in FIG. 12, acceleration occurs in the opposite direction to the direction originally intended to be obtained at the start of a turn, resulting in poor response of the aircraft.

In any of the steering methods, as shown in FIG. 8, after the flight starts, the center of gravity moves forward as the rocket motor burns, and the distance to the point of gravity of the aircraft becomes longer. This increases the stability and reduces the responsiveness of the two aircraft.

This invention was made to solve this problem, and after the rocket motor burns, a part of the two aircraft is separated, and the two
This project proposes a guided flying object whose performance can be improved by simultaneously steering the food control blades.

[Means to solve the problem]

According to this invention. The means for improving the performance of a guided flying vehicle is to detach a part of the fuselage using the fuselage's detachment mechanism after the rocket motor burns, and to attach one to each actuator of the steering wing attached to the front and rear of the main wing. By sending control signals from two control devices, these actuators are actuated simultaneously.

[Effect]

The guided flying vehicle according to the present invention has two types of steering wings: a front wing in front of the main wing and a rear wing behind the main wing. Steering is achieved by simultaneous actuation of real actuators.

When a part of the fuselage is separated, the point of force of the aircraft moves forward and becomes longer as the rocket motor burns.The distance between the center of gravity and the point of force can be shortened, reducing static stability. Improves aircraft responsiveness.

However, by simultaneously steering two types of steering blades with one control unit, it is possible to achieve both the quick response of the front wing steering system and the high trim angle of the rear wing steering system. Advantages are obtained and the performance of guided missiles is improved.

〔Example〕

FIG. 1 is a schematic diagram showing an embodiment of the present invention. The dashed line is separated from the fuselage after the rocket motor burns out. The rear end of the fuselage is shown.

The second figure is a cross-sectional view showing the internal structure of the guided flying vehicle according to the present invention.
6a) is the actuator for the front wing (2), (6b) is the actuator for the rear wing (4), and (7) is the control load (
5) is the cable that sends the control signal to the actuator (6b), (9) is the separation mechanism at the rear end of the fuselage, and aS is the rocket motor.

FIG. 3 is a sectional view showing the internal structure of the guided flying vehicle shown in FIG. 1 during operation, and Xcp2 is the point of impact of the aircraft after the rear end of the aircraft is separated.

Δx3 is the distance between the center of gravity and the point of force after the rear end of the fuselage is separated. The broken line indicates the rear end of the fuselage that will be separated from the fuselage after the rocket motor burns.

FIGS. 4 and 5 are schematic diagrams showing an example of the operation of the guided flying object shown in FIG. 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 (F), (8&) is the X axis, and (81)J 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, the white arrow indicates the direction of airflow.

In a statically stable guided flying vehicle configured as described above, the point of impact of the aircraft remains constant regardless of the combustion of the rocket motor α, as shown in Figure 2. Accordingly, the distance between the double center of gravity and the point of force is Δ
It changes in the direction of lengthening from x1 to Δx2. When this aircraft receives a signal indicating the completion of combustion from the rocket motor αG, the separation mechanism (9) at the two fuselage ends operates as shown in Fig. be separated from By separating the rear end of the aircraft, the aircraft's point of force moves forward to Xcp2, and the center of gravity position xrn2
The distance from Δx2 to Δx3 becomes shorter. By shortening the distance between the center of gravity and the point of impact, the static stability of the aircraft, which had been increasing due to the combustion of the rocket motor αG, decreased, and along with this, the responsiveness of the nine aircraft, which had deteriorated, improved. This effect can be obtained.

At the start of the turn Kk, as shown in FIG. 4, the actuator (6a) of the front wing (2) is operated in the direction of raising the leading edge of the front wing (2) by the control signal from the control unit 151. The actuator (6b) of the rear wing (4) is
It operates in the direction of lowering the leading edge of the At this time, the two aircraft have good aircraft response because the front wing (2) is steered, and it is possible to prevent or reduce the occurrence of acceleration in the opposite direction in the rear wing steering system, and only the front wing It has the same effect as steering.

During a steady turn, as shown in Figure 5, the actuator (6a) of the front wing (2) operates in the direction of raising the leading edge of the front wing (2) in response to a control signal from the control rack ft (51). , the actuator (6b) of the rear wing (4)
) moves in the direction of lowering the leading edge of the At this time, the clockwise moment required to balance one aircraft is obtained from both the front wing (2) and the rear g (41). It requires less equipment than when steering.

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

〔Effect of the invention〕

This invention is accomplished by separating a part of the rear end of the fuselage from the fuselage after the rocket starter has been burned.
After improving the aircraft's responsiveness.

1 for the actuators of the two types of control vanes, front wing and rear wing.
By sending control signals from two control devices to simultaneously steer both wings, it is possible to obtain the advantages of both the front wing steering method and the rear wing steering method.

[Brief explanation of the drawing]

FIG. 1 is a schematic diagram of a guided keratoplastic body according to an embodiment of the present invention, FIG. 2 is a sectional view showing the internal structure of the guided keratoplastic body according to the present invention, and FIG. 3 is a guided keratoplastic body according to the present invention. FIGS. 84 and 5 are schematic diagrams showing an example of the operation of the keratinoid according to the present invention. FIG. FIGS. 6 and 7 are schematic diagrams showing a conventional guided corneal gland, FIG. 8 is a schematic diagram showing a part of the internal structure of a conventional guided corneal gland, and FIGS. 9, 10, and 11. FIG. 1 and FIG. 12 are schematic diagrams showing an example of the operation of a conventional guiding gonioglial body. In the figure, (2) is the front wing, (3) is the main wing, and (4) is the rear wing. (5) is a control device, (61 is an actuator, (9)
is the disconnection mechanism. In each figure, the same reference numerals indicate the same or corresponding parts.

Claims (1)

[Claims] In a guided flying object that has a guidance section for guiding to a target and a control section for controlling the attitude of the aircraft, and that flies by obtaining thrust from a rocket motor, it has a main wing and a front section installed in front of the main wing. A wing, a rear wing provided behind the main wing, a control device for steering, a separation mechanism for separating a part of the rear end side of the aircraft from the aircraft, and a control signal receiving from the control device. A guided flying object, comprising: the front wing actuator and the rear wing actuator, which operate as follows.