JPH0830640B2 - missile - Google Patents

Description

Description: FIELD OF THE INVENTION This invention relates to steerable weapon devices and missiles. The present invention provides a "manpack" type infantry weapon device, especially in a small size and for single use.

Prior Art Corresponding missiles are designed to reduce the power of armored land targets and vehicles and the like, especially tanks. The payload carried by the missile is actually built as a combat payload. These payloads, which must be effective against new armor, have sufficient mass and caliber. These weapon devices have to be fired from a very narrow minimum range of a few tens of meters (Sielta, basements, buildings, etc.), and in various runs the maximum range is about 600-700 meters. Can be set between 2000 meters.

Problems to be Solved by the Invention Regardless of problems such as deviation when the missile is launched, deviation due to wind or air current, when using land-based anti-tank weapons exceeding the range of 300 meters, calculation, especially the center of gravity is used. It means that the trajectory of the missile must be modified in consideration of the action of and the movement of the target on a straight line. Due to ergonomics and weight considerations, it is not possible at the muzzle of a weapon to obtain a velocity that is fast enough to cover the required distance in a time unit of 1 second.

 The following two correction methods are possible.

First, a modified "program" that must be done on the one hand to correct the effect of the center of gravity, and on the other hand to measure the movement of one straight line target before firing and set on the weapon before firing. There is an open loop coaching fix to get. In firing order, this method means that the target's movement characteristics are measured and do not change significantly during the time the missile is in flight.

Next, there is a closed loop coaching modification that removes the off-axis missile offset between the launch position and the target until the missile completes its flight.

The open loop coaching modification requires a short flight, as the movement of the target is inferred, among other things:
This type of modification means for the missile, a muzzle and initial velocity sufficient to overcome the effects of wind and airflow.

The closed loop coach modification is only adapted to missiles given a low initial velocity.

Today, when a missile leaves the launch tube, it is built to use a large initial velocity to ensure that the missile has sufficient aerodynamic maneuverability. Regardless of the launching device (gas generator, booster, etc.), such means are limited in terms of weight, bulkiness, injection effectiveness, flame, noise, outgassing, etc., and launch in restricted locations, It requires considerable explosive loading that is incompatible with the separation of shots that can be fired without being discovered.

The suitability for firing in a limited location requires that a small amount of explosive charge be used for the firing. The corresponding device is
It may be of the well-known type such as a fuse booster or an injection device such as a very low injection speed cannon or recoilless gun.

The noise is greatly reduced, the ease of the gunner is increased, precautions are taken when the missile is fired, and the recoil of the weapon in the gunner can be reduced to almost zero, and in addition, The missile leaves the weapon at a slow speed. Therefore, the classical aerodynamic steering by taking the attack angle that produces the torque cannot use the steering force because the aerodynamic force is too small. As soon as the missile is launched to ensure a good maneuverability and a short response time of the missile that can oppose the influence of the center of gravity and wind and to attack a target that steers a short distance, the force, that is, the center of gravity It is possible to steer only by the driving force passing through. Such steering by steering forces results in a total or slight aerodynamic attack angle, rather a direct displacement of the center of gravity.

In the case of low speed missiles, such maneuvering forces can only be propelled at the root. Maintaining steering with steering force
As it is necessary to maintain the application of maneuvering force at the center of gravity of the missile, the consumption of propellant powder that produces such maneuvering force is large, which adversely affects the mass of the weapon. Furthermore, it is difficult to obtain centering adjustments as the maneuvering force must pass through the entire center of gravity accurately throughout the flight.

The object of the present invention is to correct such drawbacks. The present invention makes it possible to provide a light weapon device capable of accurately guiding a low speed missile without requiring delicate adjustment.

To this end, according to the present invention, there is provided a method for steering a missile that is launched from a launch tube at a low speed and provided with a propellant-type explosive-type propulsion device. The axial thrust and the steering force can be applied to the missile at the same time, and the center of gravity of the missile, like the action of the combustion of the propellant powder, is controlled so that the center of gravity passes through the above-mentioned line of action of the steering force when exiting the launch tube. This propulsion device is designed and supplied in such a way that the position changes, and during flight, it produces an aerodynamic attack angle to develop the torque that produces the steering force applied to the steering by the steering force. It is noted that this invention can be made.

Therefore, the present invention relates to being effective at a low initial velocity so that the missile is steered only by the maneuvering force at the time of launching the missile, and thus the change of the center of gravity sufficiently increases the flight speed. At times, the missiles launch tubes into a specially constructed flight propulsion device of the missile that can ensure that conversions are made into combined steering (a combination of steering maneuvering and aerodynamic steering). It consists of connecting the steering by the maneuvering force of the missile when exiting.

As will be seen below, it is advantageous to ensure during flight that the center of gravity of the missile moves back along the axis.

Optionally, a weapon system designed to carry out the method of the invention and having a launch tube and a missile, wherein the propulsion device and the control device are provided on the missile, on the one hand, before the missile leaves the launch tube. It is unique to have at least one integrated booster capable of imparting a low initial velocity to the missile when the missile leaves the launch tube after combustion and, on the other hand, at least one flight propulsion device steered by maneuvering forces. .

The missile is launched from the launch tube by an integrated booster that can impart longitudinal acceleration to the missile to set rotation on the missile, without causing any recoil at the launch position. The explosive burning or booster is completed in the launch tube, and the missile has a longitudinal velocity of 20 m / sec and about 5 revolutions /
Exit from this launch tube with a rotation speed of seconds.

A lateral nozzle of a flight propulsion device is suitable as a device for steering with a steering force.

These nozzles of the flight propulsion device are fired after the missile exits the launch tube. Nozzles on the sides leave the rear of the neat missile, thus more distributed during flight, making it easier to set up the necessary infrared electromagnetic or other coupling between the launch location and the missile. Try to Since these nozzles form an angle with the axis of the missile, the thrust of the nozzle along the axis has an axial component that acts on the longitudinal thrust. The control surface is a known device capable of deflecting propulsion jets, such as movable nozzles, baffles, jet shields, etc. of the type described, for example, in French Patent Nos. 1,099,901 and 1,602,885. Another known complex device in the middle uses a movable skirt with a nozzle outlet. Deflection of the jet causes deflection of the thrust vector associated with the launch tube and affects the steering force. The amount of deflection produced by the simple device is relatively small and the steering force is small with respect to the axial component of thrust.

When the missile is launched, the maneuvering force opposes gravity and allows the missile to have good maneuverability, thereby
It produces a doubled value of lateral load factor and produces a sufficient axial component of thrust for the flight propulsion system. Missiles are continuously accelerated.

The speed of the missile depends on when the missile exits the launch tube.
Increases throughout the trajectory from 20 m / sec to high subsonic values at the end of flight. A stabilizer attached to the rear of the missile that can be used as the missile exits the launch tube ensures aerodynamic stability even at low speeds and acts to lift during combined steering.

The thrust principle of a propulsion device can be programmed as a function of time of flight by limiting the combustion state of the propellant charge or the nature of the explosive charge (eg, a two component device). Therefore,
The big burning surface for this charge at the beginning of the flight is
It can give the missile a great deal of effectiveness with regard to the load factor it can create to quickly maneuver a moving target to compensate for the effects of gravity and wind as it leaves the launch tube.

Optimization of the combustion principle in the rest of the flight makes it possible to limit the minimum mass in the propellant charge.

In accordance with the present invention, there is the advantage that during flight the center of gravity can be changed from the position imposed on the launch. Such changes in the center of gravity take special measures with regard to the mounting of the different assembly parts that make up the missile, in such a way that they are carried out in such a way that the desired implementation can be achieved.

In particular, the flight propulsion device is attached to the front of the missile and a single component combat payload is provided at the rear, and the center of gravity of the missile moves to the rear during flight, resulting in a combined steering effect. It is possible to split the payload into two or more parts on either side of the propulsion unit. At the same time, this arrangement is particularly suitable for the action of hollow charge (s) provided on the missile by specifically increasing the attack interval (distance between charge and armor at the moment of explosion).

The present invention is described in detail below with reference to the accompanying drawings without any limitation. This description is suitable for short range anti-tank weapon devices.

EXAMPLE The anti-tank missiles I, II, III of FIGS. 1, 2, and 3 are mounted at the rear and designed to give the missile a velocity of, for example, 20 m / sec when the missile is launched from the launch tube 7. Two boosters 1a, 1b, and a flight propulsion device 2 that is fired several meters after leaving the launch tube 7 to avoid the gunner's hindrance and to ensure that the missile crosses a confined void. have. The boosters 1a, 1b are of the fuel charge type, and the combustion of this propellant charge is carried out completely within the firing tube 7 in order to avoid the hindrance of the gunner. This small increase in speed is obtained by burning an equal small amount of explosive compatible with the launch of missiles in confined voids. This flight propulsion device 2 is mounted on the same plane so as to be symmetrical with respect to the longitudinal axis LL of the missile.
It has three identical nozzles 3a, 3b. These nozzles 3
The axes of a and 3b coincide with the longitudinal axis of the corresponding missile, and are inclined symmetrically with respect to this axis. Steering command (8th
(Figs. a and 8b) without, the axial components R = Q ₁ of the thrusts F ₁ and F ₂ of the two nozzles 3a and 3b.
＋ Q ₂ can be made to fly missiles, and the effects of the thrusts F ₁ of nozzles 3a and 3b on the missiles of components P ₁ and P ₂ of F _{2 become} zero due to the symmetry in the plane of nor.

In each of the nozzles 3a and 3b, the outer cylinder of the missile has an opening 8 through which gas can pass.

Each nozzle 3a, 3b has two nozzles in response to a steering command.
The jets of 3a and 3b can be deflected at the same time, and therefore the lateral thrust component M = M ₁ + M ₂ (see FIGS. 9a, 9b and 9c) which is orthogonal to the plane of the two nozzles 3a and 3b is generated in the same direction. A jet blocking or deflecting device (not shown) of known type is provided. This lateral thrust component M can drive the missile.

Also, the missile has a hollow charge type combat payload 4,
An equipment block 5 mounted at the rear of the missile to house the equipment needed to guide and steer the missile, a retractable tail 6 to ensure aerodynamic stability, and a containment tube acting as a launch ramp, ie It has a launch tube 7.

When the missiles I, II, III leave the launch tube 7 under the action of the two boosters 1a, 1b, the combustion of the boosters 1a, 1b ends, and the center of gravity G of the missile is the lateral thrust of the thrust of the flight propulsion device 2. Lie according to the plane containing the component M. The missile is then only steered by thrust.

In the case of the missile II shown in FIG. 2, the charge, that is, the payload 4 is provided at the front part, and the main propulsion device 2 is divided into two propulsion devices 2a and 2b provided at both sides of the center of gravity G, during flight. Combustion is performed in the same state such that the center of gravity G is fixed to. This is due to the curve B in FIG. 4 showing the change in the distance l separating the front of the missiles I, II, III from the center of gravity G as a function of time T shown in FIG. 4 so that the axis of time extends from right to left. Shown.

In the missile III of FIG. 3, the payload 4 is provided at the front as in FIG. 2, but the propulsion device 2 is an integrated device and faces the payload 4 with respect to the center of gravity G. In addition, when the propellant 2 charge is used, the center of gravity G moves towards the front of the missile III and the distance l decreases.
This is shown by curve C in FIG.

Conversely, referring to FIG. 1, in the missile I described by the present invention, the flight propulsion device 2 is attached to the front part, while the charge, that is, the payload 4 is the main charge provided in the rear part, that is, the payload 4. It is divided into a payload 4a and an auxiliary charge provided in the front portion, that is, a payload 4b (which can be omitted). Therefore, when the gunpowder of the flight propulsion device 2 burns, the center of gravity G of the missile I moves to the rear part, and the lateral moment M of the thrust of the propulsion device 2 causes an attack angle, that is, a torque that causes a lift, by each of the control commands given. . Therefore, the lateral moment M and the lift, respectively generated, produce an added steering effect. In FIG. 4, the curve A
Indicates an increase in the distance l as the flight time increases.

In FIG. 5, curves D, E and F show the maneuverability of missiles I, II and III as a function of flight time. For this reason, the load coefficient n is shown on the vertical axis and the time T is shown on the horizontal axis.
This figure shows that the maneuverability of the missile described by the present invention increases with time, while the missile of FIG. 3 declines to reduce maneuverability. In the missile II of FIG. 2, the maneuverability is slightly increased due to the unstable consumption of the propellant charge, and the straight line H represents an ideal state without instability. That is, as described above, although the nozzles 3a and 3b are at the fixed positions, the centers of gravity of the missiles I and III move with time as the propellant charge 2 is consumed. Therefore, as time passes, the thrusts F1 and F2 are applied at a fixed position, but the distance between the fixed position where the thrust is applied and the movable center of gravity changes continuously. , The center of gravity moves. Therefore, the components of thrust change and the maneuvering performance of each missile changes with time.

Fig. 6 shows the velocity V (m / sec) of the missile I and the time T
The distance X traveled by the missile as a function of (seconds)
(M) is shown as an example. In this example, the velocity V increases linearly and the distance X increases as a function of time squared.

It is noted that it is suitable to divide the hollow charge type combat payload into two parts, one payload 4a part in front of the propulsion device and the other payload 4b part in the rear. In fact, for the same missile length, this structure is much larger in the main aft charge, ie payload 4b, than in the classical design as in Figures 2 and 3 where the main charge is in the front. It can be made to have a range of action (distance a between the charge and the armor at the moment of the explosion), which makes the charge more effective, ie the penetration distance b into the armor plate Greatly promotes (see Figure 7a).
Curve J in FIG. 7b shows that the penetration distance b has a maximum value as a function of the ratio a / d. Here, a is the distance in the above-mentioned working range, and d is the diameter of the missile (Fig. 7a). On curve J, the missile I described by this invention corresponds to a more preferred point K than the point N corresponding to missile III in FIG.

An embodiment of a missile has been described above which results in automatic rolling about the axis LL in such a way that two nozzles for steering by force are sufficient. If the missile is rolling-stabilized, it will have at least four lateral devices for force steering.

As shown in FIG. 1, the containment tube, or launch tube 7, can exhibit a diameter greater than the rear diameter, and the front portion of the missile having a smaller diameter than the rear portion is secured by an annular shim 9.

[Brief description of drawings]

FIG. 1 is an axial cross-sectional schematic view of a weapon device constructed in accordance with the present invention consisting of a missile in a launch tube, which uses simple steering by force at launch but at the end of flight due to aerodynamic effects. Featuring the unbalanced weight principle associated with the consumption or consumption of consumable products that enhances maneuverability, Figure 2 exclusively uses force steering throughout the flight with a constant missile center of gravity A similar cross-sectional schematic view of a missile in a launch tube, FIG. 3 shows exclusively the steering by force in launch where the unbalanced weight principle associated with consumption of consumable product leads to a reduction in missile maneuverability at the end of flight. A similar schematic cross-section of the missile used, FIG. 4 is a curve characterizing the law of change of the position of the center of gravity during a missile flight of a missile of the type described above, shown schematically in FIGS. A curve is given to show that the features can be implemented rather than a qualitative comparison, and FIG. 5 shows a curve representing the change in maneuverability of the missile of FIGS. Curves are given to allow features rather than comparisons, Figure 6 shows the curves of speed and distance changes as a function of time for the missile of Figure 1, and Figure 7a is for a hollow-charged combat payload. FIG. 7b shows the effect of the working range distance, which is characterized by the penetration distance obtained with the semi-infinite steel plate, FIG. 7b shows the penetration characteristics of the missile shown in FIGS. 1 and 3, and FIGS. 8a and 8b show FIG. Figure 9a, 9b, 9c schematically show the decomposition of the thrust vector of the nozzle of the missile without steering command, and Figs. 9a, 9b and 9c show the decomposition of the thrust vector of the nozzle of the missile with steering command. It is a rear view, a plan view, and a side view. In the figure, 1a, 1b: booster,
2: Flight propulsion device, 3a, 3b: Nozzle, 4a, 4b: Charge, that is, payload, 5: Equipment block, 6: Tail, 7: Launch tube, 8:
Aperture, 9: Annular shim.

 ─────────────────────────────────────────────────── ─── Continuation of the front page (72) Inventor Michel Murray France, 91190 Gif-sur-Yvette, Lieu de la Grayley 84 (72) Inventor Robert Weiyan France, 75007 Paris, Lieu De Sable 88 (72) Inventor Jean Ansaldi France, 78470 Saint-Remy-les-Chevrouse, Alle des Solmes 16

Claims (11)

[Claims] 1. A combat payload and a propulsion fuel load type propulsion device, which is provided at the rear of the missile, burns before the missile exits the launch tube, and when the missile exits the launch tube, 20 m The first propulsion device that gives a low initial velocity of about 1 / second and the missile can be steered by force, and it is provided to simultaneously give the missile a steering force for steering by axial thrust and force during flight of the missile. A second propulsion device having a transverse nozzle, wherein the combat payload and the first and second propulsion devices traverse the longitudinal axis of the missile when the missile exits the launch tube after combustion of the first propulsion device. The missile is arranged in the plane and has a steering force, and the position of the center of gravity of the missile is the longitudinal axis of the missile during flight of the missile and during combustion of the second propulsion device. A missile fired from a launch tube that is capable of producing a torque that travels backwards along which causes another steering force to be added to the steering force provided for steering by the jet force. 2. The missile according to claim 1, wherein the second propulsion device is provided at a front portion of the missile, and at least a part of the combat payload is provided at a rear portion of the missile. 3. The missile according to claim 2, wherein the combat payload has a main charge portion provided at a rear portion of the missile and an auxiliary charge portion provided at a front portion of the missile. 4. The missile according to claim 2, wherein the combat payload is provided in front of the first propulsion device. 5. The missile according to claim 3, wherein the main charge portion is provided in front of the first propulsion device, and the auxiliary charge portion is provided in front of the second propulsion device. 6. The missile according to claim 1, wherein the combat payload has at least a main charge portion, and the main charge portion is provided in front of the first propulsion device. 7. The missile according to claim 6, wherein the combat payload has an auxiliary charging portion, and the auxiliary charging portion is provided in front of the second propulsion device. 8. The missile according to claim 1, wherein the second propulsion device is provided to burn symmetrically with respect to the longitudinal axis of the missile. 9. The missile according to claim 1, wherein the missile is rotationally stabilized and the second propulsion device has at least four lateral nozzles steerable by force. 10. A missile according to claim 1 wherein the missile automatically pivots about a longitudinal axis and the second propulsion device has only two lateral nozzles which can be steered by force. 11. A missile according to claim 1 having a launch tube for launching the missile.