EP0081421B1 - Terminal guidance method and guided missile using it - Google Patents

Abstract

A guidance method is provided for the terminal portion of the trajectory of a guided missile having a sensor and comprising two sections coupled together by a central shaft and free to rotate with respect to one another about the longitudinal axis of the missile; one section comprises a drive means for controlling the roll attitude of this section and a gas generator which feeds a nozzle for providing a transverse throat force and the other section has a stabilizing tail unit formed by a set of fins able to be opened out.

Description

The invention relates to guided projectiles and relates, more specifically, to a method of guiding a missile, applicable during the terminal portion of the flight path; it also relates to a guided missile operating according to this guidance method.

There is a demand for AIR-SOL missiles capable of stopping, at relatively large distances, the threat posed by land formations constituted, in particular, by motorized vehicles such as armored vehicles advancing in groups on the ground. These armored vehicles, by their nature, radiate thermal energy and, therefore, constitute potential targets which can be detected and located by a missile fitted, for example, with an electrooptical sensor E.0 operating in the LR band of the electromagnetic spectrum. In addition, the missile may be provided with a military charge capable of perforating the protective armor of armored vehicles. It is possible to direct the firing of such a missile towards a group of armored vehicles; however, the problem remains of providing, during the terminal portion of the descent to the ground trajectory, the trajectory corrections necessary to achieve an impact of the missile on one of the vehicles detected by the E.O.

Already known from US-A-3 843 076 is a projectile comprising guide means which make it possible, in the terminal phase of the trajectory, to correct the possible error between the direction of a target and the direction of impact of the projectile on the ground, in free fall. To this end, the base of this projectile of the prior art is equipped with a set of fins which imparts to the body of the projectile an autorotation movement of substantially constant angular speed, around its longitudinal axis. In the head of the projectile is arranged an electro-optical sensor (E.O) and, finally, in the middle part of the body, a lateral impeller can provide a predetermined thrust force whose direction is normal to the velocity vector of the projectile. The E.O. sensor consists of a plurality of photodetector cells arranged in a ring in a plane perpendicular to the axis of the projectile, in order to provide a hollow conical field of vision. Thus, the surface of the ground covered by the field of vision of the sensor E.0 gradually decreases as a function of the decreasing altitude of the trajectory. When the target enters the field of vision of the sensor, its image falls on one of the photodetector cells, which determines, in polar coordinates, the position of the target relative to the orientation of the impeller. The output signal from the sensor E.0 is used to provide a trigger order to the lateral impeller at the moment when the orientation of the latter is opposite to the direction of the detected target.

Other types of projectiles are described in the following patent documents. FR-A-2 230 958 describes, in a summary manner, a method of attack, from a submarine while diving, of objectives flying at low altitude, using a missile provided with 'a research head.

According to the method, is started vertica _J ement and the missile self-rotation that the search head is initially locked in a position substantially perpendicular to the axis of the missile. Because of this position and the rotation, the entire horizon is scanned. As soon as a target is detected, the search head is unlocked in order to pursue the target then simultaneously the rocker is rocked in the direction of the target being pursued.

Furthermore, American patent US-A-2,520,433 describes a missile provided with a sensor / sensor sensitive to the energy radiated by a potential target.

The missile includes first and second main sections mutually coupled and free to rotate relative to each other about the longitudinal axis of the missile body. The first section contains a sensor, a gas generator which, of course, feeds a side nozzle to provide a transverse thrust force. In the first section there is an amplifier 80 which gives a generating member 71, 72 the piloting order to trigger the transverse thrust force in order to vary the roll attitude of the missile body.

In addition, German patent DE-B-1 092 313 describes a method of guiding a missile, during the terminal portion of its trajectory, more precisely during the pursuit of a target which has been detected. This patent also describes said missile. This comprises a front section, in rotation (at a first speed) around a first axis, and a rear section, coupled to the front section, in rotation (at a second speed) around a second axis, the first axis rotating itself around the second axis making a constant angle with the latter. These respective rotations are provided by a drive member. The front section includes a sensor sensitive to the energy radiated by a potential target, and fins which make a variable and adjustable angle with said front section, a variation of this angle creating a transverse thrust force. The rear section is equipped with a stabilizing stabilizer. The missile further comprises a generator of piloting orders, to which the sensor delivers signals (which depend on the respective directions of said second axis and of a missile / target line of sight), this generator then transmitting an order to said motor member piloting which modifies the angle between said second axis and said line of sight in order to orient the missile at the target.

These prior art projectiles of relatively simple construction do not make it possible to achieve the desired degree of efficiency and, in particular, to achieve a probable impact on the target. To achieve this goal, the proposed guidance method uses a target tracking sensor which measures the rotation of the missile / target line of sight.

• - lesdits organe moteur et arbre central permettant une rotation relative entre les deux sections du missile autour d'un même axe, cet axe étant l'axe longitudinal du corps du missile,
• - la première section ayant des moyens pour créer une force de poussée transversale normale à la direction de la vitesse de déplacement du missile,
• - la première section du missile étant munie d'un senseur sensible à l'énergie rayonnée par une cible potentielle.

According to the invention, the guidance method, during the terminal portion of its trajectory, applies to a missile having a first section coupled by a central shaft with a second section, the first section comprising a driving member:

• - said motor member and central shaft allowing relative rotation between the two sections of the missile about the same axis, this axis being the longitudinal axis of the missile body,
• the first section having means for creating a transverse thrust force normal to the direction of the speed of movement of the missile,
• the first section of the missile being provided with a sensor sensitive to the energy radiated by a potential target.

• - immobiliser le faisceau de réception du senseur sur l'axe longitudinal du missile;
• - imprimer au corps du missile une rotation de vitesse angulaire de roulis déterminée autour de l'axe longitudinal du missile;
• - créer ladite force de poussée transversale et maintenir la même force pendant le reste de la phase terminale, pour imprimer au corps du missile un mouvement hélicoïdal, de sorte qu'il effectue un balayage en spirale dudit faisceau, et,
• - détecter l'image d'une cible éventuellement captée par le faisceau de réception du senseur;

The method includes, for the search for the target, the following sequences:

• - immobilize the sensor reception beam on the longitudinal axis of the missile;
• - impart to the missile body a rotation angular speed of roll determined around the longitudinal axis of the missile;
• - create said transverse thrust force and maintain the same force during the rest of the terminal phase, to impart a helical movement to the body of the missile, so that it performs a spiral scan of said beam, and,
• - detect the image of a target possibly captured by the reception beam of the sensor;

• - libérer le faisceau de réception du senseur;
• - maintenir l'axe de ce faisceau pointé sur la cible détectée, ledit axe du faisceau formant une ligne visée missile/cible;
• - mesurer la vitesse de la rotation de la ligne de visée,
• - élaborer un ordre de pilotage proportionnel à la grandeur mesurée de la vitesse de la ladite rotation de la ligne de visée,
• - appliquer cet ordre de pilotage à l'organe moteur de sorte qu'il fasse tourner la première section relativement à la seconde section du missile, ceci dans le but de modifier l'angle de roulis de la première section, pour annuler la rotation de la ligne de visée par action de la force transversale, de sorte que soit corrigée la direction de la trajectoire du missile vers la cible.

the method comprising, for tracking the target after detection, the following steps:

• - release the sensor reception beam;
• - Maintain the axis of this beam pointed at the detected target, said axis of the beam forming a missile / target line of sight;
• - measure the speed of the line of sight rotation,
• - develop a piloting order proportional to the measured magnitude of the speed of said rotation of the line of sight,
• - apply this order of piloting to the engine member so that it makes turn the first section relative to the second section of the missile, this in order to modify the roll angle of the first section, to cancel the rotation of the line of sight by action of the transverse force, so that the direction of the trajectory of the missile towards the target is corrected.

• - la section avant comportant en outre un senseur sensible à l'énergie rayonnée par une cible potentielle et des moyens pour fournir une force de poussée transversale,
• - l'organe moteur comportant une entrée de commande connectée à un générateur d'ordres de pilotage,
• - le missile comportant à sa base un empennage stabilisateur en forme d'ailettes,

The invention also relates to a guided missile, comprising a front section coupled by a central shaft with a rear section, the front section comprising a drive member, said drive member and central shaft allowing relative rotation between the two sections of the missile,

• the front section further comprising a sensor sensitive to the energy radiated by a potential target and means for providing a transverse thrust force,
• - the drive member comprising a control input connected to a control command generator,
• - the missile comprising at its base a stabilizing stabilizer in the form of fins,

characterized in that:

• the two sections of the missile are in relative rotation about the same axis, this axis being the longitudinal axis of the missile body,
• - the sensor is provided, on the one hand with a locking device for immobilizing during the search phase the sensor receiving beam along the longitudinal axis of the missile, and on the other hand with means for unlocking the sensor by detecting the target and to maintain the receiving beam axis pointed at the detected target, said beam axis forming a missile / target line of sight, a piloting order being created in response to the speed of rotation of the line of sight,
• the drive member comprises a first member secured to the structure of the front section and a second member physically coupled to the rear section,
• said control input is connected to the pilot generator via an amplifier,
• - the drive unit modifies, according to the piloting order, during the tracking phase, the roll angle of the front section,
• the means for supplying the transverse thrust consist of a gas generator supplying a lateral nozzle,
• - the stabilization fins are deployable.

Another object of the invention consists in conferring on the missile an initial speed of displacement determined on its trajectory and in maintaining it substantially constant along the trajectory.

Another object of the invention is to vary the angular speed of autorotation of the missile body along its terminal trajectory. In addition, the second member of the engine member is coupled to the rear section of the missile by a central shaft.

According to another object of the invention, the rear section of the missile comprises a compartment for housing a releasable braking parachute intended to reduce the ballistic speed of the missile over the portion of the trajectory preceding the terminal phase.

• - la figure 1 représente un projectile guidé de l'art antérieur,
• - la figure 2 représente le mode de réalisation du senseur électrooptique du projectile de l'art antérieur,
• - la figure 3, sous une forme schématique simplifiée, représente un missile guidé comprenant les moyens nécessaires à la méthode de guidage selon l'invention,
• - la figure 4 représente une vue en coupe transverse du missile guidé de la figure 3,
• - la figure 5 est un diagramme plan d'axes x, z liés au sol indiquant les principaux paramètres qui déterminent l'étendue du sol balayé par le faisceau du senseur,
• - la figure 6 est un diagramme selon un triédre x, y, z lié au sol illustrant la méthode de recherche d'une cible potentielle,
• - la figure 7 représente une vue détaillée d'une portion de la trajectoire du missile,
• - la figure 8 est un diagramme simplifié représentant une variante de la trajectoire de recherche,
  ta figure 9 illustre la loi d'accéleration conferée au missile en fonction de la vitesse de rotation de la ligne de visée missile/cible,
• - la figure 10 illustre la loi de contrôle de l'attitude de roulis du corps du missile en fonction de la vitesse de rotation de la ligne de visée missile/cible,
• - la figure 11 est une coupe longitudinale d'un missile guidé selon l'invention,
• - la figure 12 représente, en vue éclatée, les éléments d'un moteur-couple électrique,
• - la figure 13 représente un mode de réalisation de l'empennage stabilisateur,
• - la figure 14 illustre une application du missile guidé à la destruction d'un groupement de véhicules terrestres,
• - la figure 15 est une vue éclatée du compartiment d'emport d'un projectile porteur contenant une pluralité de missiles,
• - la figure 16 est une vue en coupe du projectile porteur montrant la disposition relative des missiles guidés dans le compartiment d'emport.
• - la figure 17 est un diagramme des composantes du vecteur rotation de la ligne de visée missile-cible dans un trièdre absolu et dans le trièdre missile.
• - la figure 18 représente, sous la forme d'un bloc diagramme, les éléments de la boucle d'asservissement en poursuite du missile.

The characteristics and advantages of the invention will emerge from the detailed description which will follow, made with reference to the appended drawings which illustrate the guidance method and an embodiment of the guided missile; on these drawings:

• FIG. 1 represents a guided projectile of the prior art,
• FIG. 2 represents the embodiment of the electrooptical sensor of the prior art projectile,
• FIG. 3, in a simplified schematic form, represents a guided missile comprising the means necessary for the guidance method according to the invention,
• FIG. 4 represents a cross-sectional view of the guided missile of FIG. 3,
• FIG. 5 is a plane diagram of axes x, z linked to the ground indicating the main parameters which determine the extent of the ground swept by the beam of the sensor,
• FIG. 6 is a diagram according to a trihedron x, y, z linked to the ground illustrating the method of searching for a potential target,
• FIG. 7 represents a detailed view of a portion of the trajectory of the missile,
• FIG. 8 is a simplified diagram representing a variant of the search trajectory,
  FIG. 9 illustrates the acceleration law imparted to the missile as a function of the speed of rotation of the missile / target line of sight,
• FIG. 10 illustrates the law for controlling the roll attitude of the missile body as a function of the speed of rotation of the missile / target line of sight,
• FIG. 11 is a longitudinal section of a guided missile according to the invention,
• FIG. 12 represents, in an exploded view, the elements of an electric torque motor,
• FIG. 13 represents an embodiment of the stabilizing stabilizer,
• FIG. 14 illustrates an application of the guided missile to the destruction of a group of land vehicles,
• FIG. 15 is an exploded view of the compartment for carrying a carrying projectile containing a plurality of missiles,
• - Figure 16 is a sectional view of the carrier projectile showing the relative arrangement of guided missiles in the carrying compartment.
• FIG. 17 is a diagram of the components of the rotation vector of the missile-target line of sight in an absolute trihedron and in the missile trihedron.
• - Figure 18 shows, in the form of a block diagram, the elements of the servo loop in pursuit of the missile.

FIG. 1 represents, in a simplified form, the projectile of the prior art according to US-A-3 843 076 as well as the corresponding terminal guidance method. The projectile 1 is equipped with a set of fins 2, the configuration of which makes it possible to print on the body of this projectile an angular speed of autorotation ω _r around its longitudinal axis X carrying the displacement speed vector V of the projectile on its path. In free fall, the trajectory of the projectile is inclined by an angle θ _t and this projectile strikes the ground at a point 4 angularly offset by an angle θ _c of a potential target 6.

In order to modify the trajectory of the projectile, it is equipped with a lateral impeller 3 and an electrooptical sensor 5 which provides a trigger signal for this impeller, this trigger signal resulting from the measurement of the angle error 8 _c . It follows that the velocity vector V of the projectile is modified by an amount V _c to provide a resulting velocity vector V _r offset by the angle θ _c of the velocity vector V to achieve the impact of the projectile on the target.

FIG. 2 represents the embodiment of the electrooptical sensor 5 carried by the projectile 1 described in FIG. 1. This sensor E.0 is a sensor essentially consisting of a plurality of photoconductive elements 7 arranged in a ring in a plane orthogonal to l longitudinal axis X of the projectile body to provide a predetermined hollow conical field of view of angular width. When the image 8 of the target 6 falls on one of the photoconductive elements 7 such as the element 7 _j , the magnitude of the relative angle A between the direction of the impeller 3 and the photoconductive element 7 is measured by the sensor E.0 and supplied to a calculation circuit which determines the instant of triggering of the impeller 3 corresponding to the passage of the latter in the direction of the detected target.

auxquelles correspondent l'accélération normale y_N donnée par la relation suivante

et l'accélération longitudinale γ_L donnée par la relation suivante:

où M est la masse du missile et g la grandeur du champ de pesanteur terrestre.FIG. 3 represents, in a simplified schematic form, a guided missile 10 which includes specific means of the terminal guidance method according to the invention. This missile comprises: a sensor 11, sensitive to the energy radiated by a potential target, located in the head of the missile, a means 12 for providing a transverse thrust P _o passing through the center of gravity G of the missile and a means 13 for control the roll angle 0 (Fig. 4) of the missile body 10 around its longitudinal axis X. The sensor is provided with a locking means making it possible to immobilize its beam on the longitudinal axis X, means of detection of the possible presence of a target intercepted by this beam and of angular tracking means for measuring the speed of rotation η of the target / missile line (LOS). The means 12 for providing a transverse thrust P _o comprises a combustion chamber which supplies a lateral nozzle whose thrust direction is inclined, by an angle n / 2 - a, on the longitudinal axis X of the missile; it follows that the transverse components F _N and longitudinal F _L of the force F applied to the missile are given by the following relationships:

to which correspond the normal acceleration y _N given by the following relation

and the longitudinal acceleration γ _L given by the following relation:

where M is the mass of the missile and g the magnitude of the earth's gravity field.

FIG. 4 represents a section of the missile 10, of axes X, Y, and Z; and shows the components Fy and F _Z of the normal force F _N as a function of the roll angle 0 of the missile body around its longitudinal axis X. These components Fy and F _z are given by the following relationships:

The missile body can rotate in both directions, relative to the X axis with an instantaneous angular speed or autorotation speed 0. The quantities 0 and 0 can be measured on board the missile and used respectively to control the angle roll 0 and autorotation speed 0 of the missile body.

_o sur l'horizontale. Le corps du missile étant l'objet d'une vitesse d'autorotation 0 autour de son axe longitudinal X et le faisceau 14 du senseur E.0 étant immobilisé sur cet axe longitudinal X, il en résulte que le faisceau 14 décrit en fonction du temps un cône creux d'axe G.I. dont les demi- ouvertures externe et interne ont pour valeurs respectives (0 + ε) et (0 - e). L'altitude R_h du missile au-dessus du sol diminuant proportionnellement au temps, l'axe 15 du faisceau 14 décrit sur le sol, en fonction du temps, une spirale convergente de rayon R_s centrée sur le point I. L'étendue de la surface du sol balayée par le faisceau 14 est un cercle lorsque l'angle de descente est égal à 90° et une ellipse de faible exentricité lorsque la valeur de cet angle θ_o reste élevée, 60 à 70° par exemple.FIG. 5 is a plane diagram of axis x, z linked to the ground on which are indicated the main parameters which determine the extent of the ground swept by the reception beam 14 of the sensor EO carried by the missile 10 described above. The center of gravity G of the missile is animated by a speed of movement V directed along the longitudinal axis X of the body of the missile and it is subjected to a system of forces comprising: the normal force F _N to which corresponds an acceleration γ _N normal to the speed vector V, the longitudinal force F _L to which corresponds an acceleration γ _L directed along the longitudinal axis X and the earth gravity force to which corresponds the acceleration vector g directed along the vertical of the place. The sensor beam 14 has a relatively narrow half-opening angular field ε, a few degrees for example. The line GI of the missile's descent trajectory is tilted at an angle

_o on the horizontal. The missile body being the subject of an autorotation speed 0 around its longitudinal axis X and the beam 14 of the sensor E.0 being immobilized on this longitudinal axis X, it follows that the beam 14 described according to the time a hollow cone of axis GI whose external and internal half-openings have the respective values (0 + ε) and (0 - e). The altitude R _h of the missile above the ground decreases in proportion to time, the axis 15 of the beam 14 describes on the ground, as a function of time, a converging spiral of radius R _s centered on point I. The extent of the surface of the ground swept by the beam 14 is a circle when the angle of descent is equal to 90 ° and an ellipse of low eccentricity when the value of this angle θ _o remains high, 60 to 70 ° for example.

du missile autour de son axe longitudinal X est maintenue constante ainsi que la vitesse V du missile en négligeant la force de résistance de l'air et en considérant que la force d'accélération γ_L longitudinale produite par la tuyère du missile et la force de pesanteur g sont de valeurs égales et opposées. La trajectoire S du centre de gravité G du missile décrit une hélice portée par un cylindre 16 d'axe z vertical passant sensiblement par le point 1 et le rayon de ce cylindre a une grandeur r. L'étendue A_s de la surface du sol balayée par le faisceau 14 du senseur E.0 est donnée par la formule suivante:

FIG. 6 is a diagram in a trihedron x, y, z linked to the ground which illustrates the method of search for a target by the missile described previously, in a particular case corresponding to a descent angle θ _o equal to 90 °. We consider here the case where the speed of rotation

of the missile around its longitudinal axis X is kept constant as well as the speed V of the missile by neglecting the resistance force of the air and considering that the longitudinal acceleration force γ _L produced by the nozzle of the missile and the force of gravity g are of equal and opposite values. The trajectory S of the center of gravity G of the missile describes a propeller carried by a cylinder 16 of vertical z axis passing substantially through the point 1 and the radius of this cylinder has a magnitude r. The extent A _s of the ground surface swept by the beam 14 of the sensor E.0 is given by the following formula:

The surface of the ground ΔA _s intercepted by the optical beam is an ellipse whose magnitudes of the axes ΔR _s and ΔR ' _s are given respectively by the following relationships:

The oblique distance R _d , between the missile and the surface ΔA _s of the ground intercepted by the beam of the sensor EO, is given by the following relation:

The horizontal distance R _s between point I and the center of the surface ΔA _s is given by the following relation:

In this figure 6, a target c animated with a speed V _c and distant by a value R _e from point I has also been indicated. To ensure a high probability of detection of a target such as c, the angular speed Ω around the vertical axis 7 of the beam 14 of the sensor E.0 must be determined to obtain a certain degree of overlap of the successive scanning frames.

où Ω est la vitesse de rotation angulaire du faisceau.The transit time T _D of the optical beam on a target c is given by the following relation:

where Ω is the angular speed of rotation of the beam.

FIG. 7 represents a detailed view of a portion of the trajectory S of the missile 10 shown in the previous figure. The speed vector V of the missile originates from the point G representing the center of gravity of the missile, this velocity vector V is contained in a plane P tangent to a generator of a cylinder 16 carrying the point G. The components of the velocity vector V are the vertical component V _h and the orthogonal component V _t given by the following relationships:

avec

The speed component V _t is tangent to the circle with center 0 and radius r. General relationships of dynamics

with

By combining the preceding relations, the value of the angle of inclination θ of the speed vector V of the missile is obtained, with respect to the generator GI of the cylinder.

H correspondant à la distance séparant le centre de gravité G du centre de la surface ΔA_s

en considérant que les valeurs des angles ε et θ ont des valeurs toujours faibles.FIG. 8 is a simplified diagram representing a variant of the method of searching for a target on the ground. According to this variant, the angular speed 0 of roll of the missile, around its longitudinal axis X, is varied as a function of the altitude R _h of the missile above the ground. The preceding formulas giving the values of the width ΔR _s of the successive scanning frames and the angle of inclination θ of the speed vector V of the missile can be rewritten in an approximate form:

H corresponding to the distance separating the center of gravity G from the center of the surface ΔA _s

considering that the values of the angles ε and θ have always small values.

It can be shown that if the adjacent scanning frames of the EO sensor beam overlap with an overlap factor of 50%, we have the following relationship:

As a result, the trajectory S of the center of gravity G of the missile is inscribed on the surface of a cone.

est sensiblement satisfaite, le senseur E.0 détecte l'image de la cible. A partir de cet instant, le senseur E.0 fournit les signaux de sortie suivants: un premier signal de sortie indiquant la présence d'une cible dans le faisceau 14 et un second signal de sortie proportionnel à la vitesse de rotation ṅ de la ligne de visée missile/cible. Le premier signal de sortie est utilisé pour libérer le faisceau du senseur optique et autoriser la poursuite angulaire du senseur sur l'image de la cible; le second signal de sortie, une fois la poursuite angulaire assurée, est fourni à un moyen de calcul pour contrôler l'angle de roulis 0 de la section avant (voir page 13 fig. 11) du corps du missile et, par voie de conséquence, de piloter le missile en direction.We have just analyzed in detail the initial portion of the terminal trajectory of the missile corresponding to the search phase for a possible target located in an area A _s of the ground centered on the descent axis of the missile. In what follows, the final portion of the missile trajectory will be described corresponding to the acquisition of the image of the target by the sensor and, subsequently, to the piloting of the missile to achieve an impact on the detected target. Referring again to FIGS. 6 and 7, it can be seen that, when the plane P, in its rotational movement with respect to the vertical axis z, passes, at a given instant, in the vicinity of the point C corresponding to the position d 'a target and that the following relationship:

is appreciably satisfied, the sensor E.0 detects the image of the target. From this instant, the sensor E.0 provides the following output signals: a first output signal indicating the presence of a target in the beam 14 and a second output signal proportional to the rotation speed ṅ of the line missile / target sighting. The first output signal is used to release the beam from the optical sensor and allow angular tracking of the sensor on the target image; the second output signal, once the angular tracking is assured, is supplied to a calculation means for controlling the roll angle 0 of the front section (see page 13 fig. 11) of the missile body and, consequently , to pilot the missile in direction.

FIG. 9 is a diagram which represents the acceleration as a function of the speed of rotation ṅ of the missile / target line of sight, _YN being the acceleration corresponding to the force F _N of normal thrust to the speed vector V passing through the longitudinal axis X of the missile and A0 the orientation angle of this thrust force corresponding to the roll angle 0 in FIG. 4.

qui correspond à une loi de navigation proportionnelle de gain A comportant un biais η_o, γ_η représentant l'accélération en rotation de la ligne de visée missile/cible. Si, à titre d'exemple, on fait correspondre à ce biais l'accélération

ce qui a l'avantage de donner une marge de manoeuvrabilité égale de part et d'autre de la grandeur ṅ_o donnée par la relation suivante:

The equation for the missile piloting law is of the form:

which corresponds to a proportional gain navigation law A comprising a bias η _o , γ _η representing the acceleration in rotation of the missile / target line of sight. If, for example, we make this acceleration correspond

which has the advantage of giving an equal margin of maneuverability on both sides of the magnitude ṅ _o given by the following relation:

puisque les termes ṅ_o et V de l'équation de la loi de guidage sont des constantes.Consequently, the piloting input signal is proportional to the quantity ṅ and the response is the quantity A0 of the orientation of the thrust force F _N relative to the direction of the rotation vector ṅ such that

since the terms ṅ _o and V of the equation of the guide law are constants.

Figures 9 and 10 shown opposite, illustrate the laws of acceleration y and the roll steering angle A0 of the missile as a function of the module of the rotation vector ṅ.

FIG. 17 is a diagram showing the components of the rotation vector il in an absolute trihedron U, V and in the missile trihedron Y, Z referenced to the direction of the pilot nozzle.

FIG. 18 represents, in the form of a block diagram, the servo-control loop in pursuit of the missile which comprises the following elements: the guidance sensor 100 which delivers the components ṅ _y and ṅ _z of the vector of the speed of rotation of the missile-target line of sight, these two components are supplied to a resolver device 110 and an operator 120 which develops the module of the rotation vector | ṅ |, this rotation vector | ṅ | is applied to an operator 130 to supply an output signal A0 in accordance with the law of guidance shown in FIG. 10 and by means of a servo motor 140, turns the resolver 110 by an equivalent angle; finally, the output signal V ε is applied to the roll control means 150 of the front section of the missile body corresponding to the means 13 in FIG. 3, and to the motor member 24 of FIG. 11.

dans laquelle V_R est la vitesse relative et R_d la distance restante missile-cible. Il en résulte que la composante d'accélération y_N assure une navigation proportionnelle biaisée et la composante d'accélération y_T engendre une trajectoire spirale mais n'a pas d'effet sur la convergence du guidage sur la cible.The crossed component of the acceleration y _T = γ _N sin A0 generates a spiral movement of the intercept trajectory of the missile. The angular velocity 0 of roll of the front section of the missile body is then given by the following relation:

in which V _R is the relative speed and R _d the remaining missile-target distance. It follows that the acceleration component y _N provides biased proportional navigation and the acceleration component y _T generates a spiral trajectory but has no effect on the convergence of the guidance on the target.

The guidance method which has just been described can be applied to a guide missile of moderate caliber, for example of the order of 100 mm, and the magnitudes of the main parameters listed above may, for information, be around the following values: displacement speed V of the missile on its trajectory of the order of 50 ms ^-1 , descent angle θ _o between 60 and 90 °, tilt angle 0 of the missile speed vector on the axis of descent between 10 and 15 °, angular half-opening ε of the sensor beam of the order of 4 to 8 °, altitude R _h of the missile at the time of ignition of the gas generator, of the order of 500 m. For these values of the main parameters, the duration of travel of the terminal portion of the trajectory is between 10 and 15 seconds and, for a value of normal acceleration y _N of the order of 25 ms- ² , the angular speed of rotation in roll 0 is of the order of 2.5 rad.s ^-1 , the surface of the ground swept by the beam of the sensor is approximately 5.10 ⁴ m ² . All the values of these parameters can vary depending on the specific mission of the missile.

FIG. 11 is a view in longitudinal section of an embodiment of a guided missile operating in accordance with the guidance method which has just been described.

• - un senseur E.0 23 situé derrière un dôme transparent 23a,
• - un organe moteur 24 permettant de contrôler l'angle 0 de roulis de cette section avant; cet organe moteur comprenant: un premier membre 24a solidaire de la structure mécanique de cette section avant et un second membre 24b physiquement couplé à l'arbre central 21 d'accouplement des sections avant et arrière du missile,
• - un compartiment 25 rassemblant les circuits électroniques associés au senseur E.O, d'une part, et à l'organe moteur 24, d'autre part, et
• - un générateur de gaz 26 couplé à une tuyère latérale 27 dont l'orifice de sortie est situé sur la paroi latérale externe de cette section avant.

The guided missile 10 comprises two main sections: a first main section 20, called "front section" and a second main section 30 called "rear section" which are free to rotate relative to each other about the axis longitudinal X of the missile. The front and rear sections are mutually coupled by the intermediary of a central shaft 21 carried by two bearings 22a and 22b. Inside the front section 20 are arranged the following elements:

• - an E.0 sensor 23 located behind a transparent dome 23a,
• - a drive member 24 for controlling the roll angle 0 of this front section; this drive member comprising: a first member 24a secured to the mechanical structure of this front section and a second member 24b physically coupled to the central shaft 21 for coupling the front and rear sections of the missile,
• a compartment 25 bringing together the electronic circuits associated with the EO sensor, on the one hand, and with the motor member 24, on the other hand, and
• - A gas generator 26 coupled to a side nozzle 27 whose outlet orifice is located on the external side wall of this front section.

• - la charge militaire 33 du missile, et
• - un compartiment de rangement 34 d'un parachute 35 libéré sur la trajectoire du missile, puis largué en vol.

The rear section 30 of the missile, physically integral with the central coupling shaft 21, is provided at its base with a stabilizing stabilizer 31 formed by a set of deployable fins 32; in this figure, only two fins have been shown; one of the fins 32a is shown in the deployed or active position while the other fin 32b is shown in the folded or inactive position. Inside this rear section are the following elements:

• - the military charge 33 of the missile, and
• - A storage compartment 34 of a parachute 35 released on the trajectory of the missile, then dropped in flight.

Such a missile can be characterized by its following main dimensional parameters: its caliber equal to its outside diameter D _o , its overall length L _o , the span of its fins LE and its total mass M _o .

The main elements listed below will now be described. The sensor E.0 23 is a sensor sensitive, for example, to the energy of thermal origin radiated by the vehicles to be intercepted and the dome 23a is transparent to the corresponding IR radiation. This sensor E.0 comprises an optical assembly at the focal point of which a photodetector element 23c is arranged to supply a beam 14 for receiving a half-opening equal to a quantity ε, this beam being materialized by its axis 15. The assembly constitutes by the optical assembly and the photodetector element 23c is carried by a gyroscope comprising locking means (tuliping) for immobilizing the axis of the optical beam 14 on the longitudinal axis X of the missile and precession means making it possible, in the unlocked position, to orient this optical beam in the space. In addition, this sensor E.0 comprises electronic means for detecting the presence of a thermal source intercepted by the beam and means for controlling the axis of the optical beam on the missile / target line.

The drive member 24 for controlling the roll angle of the front section of the missile is a torque engine. A torque motor is a rotary multipolar electric machine which can be coupled in direct engagement with the load to be driven. This type of machine transforms electrical control signals into a mechanical torque large enough to obtain a determined degree of precision in a speed or position control system. A torque motor of the "pancake" type, by design, can be easily integrated into the structure of the missile. As shown in FIG. 12, this type of torque motor essentially comprises three elements: a stator 24a which provides a permanent magnetic field, a laminated rotor 24b, wound, secured to a blade collector 24c, and a brush holder ring 24d equipped with connections intended to receive control signals. By virtue of its mechanical characteristics, this torque motor ensures rigid coupling with the load, resulting in a high mechanical resonance frequency; due to its electrical characteristics, the intrinsic response time of a torque motor can be short and its resolution high. In addition, the delivered torque increases in proportion to the input current and is independent of the speed or the angular position. The torque being linear as a function of the input current, this type of machine is free from an operating threshold. Torque motors are marketed, in particular, by the firms ARTUS (France) and INLAND (U.S.A.). The second member 24b of the engine member, due to its connection with the tail flammed rear part of the missile, is the object of a resistant torque resulting from the combination of the inertia torque of this rear section and the aerodynamic torque supplied by the tail. The first member 24a of the drive member has a control input which is connected to an amplifier which includes corrective electrical networks. The input of this amplifier, during the search phase of a target by the sensor, receives an electrical signal resulting from the comparison of the angular velocity Ô of roll of the missile body and a set value. The angular roll speed of the missile body can be provided by a gyrometer whose sensitive axis is aligned with the longitudinal axis of the missile. The set value can be varied as a function of time, that is to say as a function of the altitude of the missile above the ground. During the piloting phase of the missile on the detected target, the input of the amplifier of the motor organ receives an electrical signal making it possible to control the roll angle of the missile body in order to cancel the rotation of the missile / target line of sight.

The empennage 31 of the missile is constituted by movable fins between a position folded against the body of the missile and an active deployed position. Taking into account the relatively low speed of movement V of the missile, it is necessary that the tail unit provides a significant aerodynamic stabilizing torque, this is obtained by fins of great elongation which are placed tangentially on the body of the missile. Figure 13 is a perspective view of the entire empennage, the fins located on the front of the figure being deleted for clarity. The body 31a a of the tail unit is an annular part provided, for example, with an internal thread 31b allowing its fixing on the base of the rear section 30 of the missile. This annular part includes a set of inclined yokes 31c regularly distributed around the periphery of the part. In these yokes, a slot 33 with parallel faces allows to embed the hinge tab 34 of the fin 32 which can pivot, by means of a pin in the holes 33a and 33b. From the mechanical point of view, the tail is supplemented, for each of the fins, by a locking device in the deployed position. This device is constituted, for example, by a spring locking mechanism 36 which actuates a stud 37, which can engage in a lateral notch provided for this purpose in the hinge tab of the fin. A detailed embodiment of this type of tail has been described in the French patent PV. n ° 53 419, deposited on March 15, 1966 and published under n ° 1 485 580. In addition to its stabilizing function, the empennage provides an aerodynamic resistant torque which is transmitted to the second member 24b of the motor member 24.

The gas generator 26 is essentially constituted by a combustion chamber inside which are arranged two blocks 26a and 26b of solid propellant. Between these two propellant blocks is located an ejection nozzle 27, the outlet opening of which opens onto the side wall of the missile body. The thrust direction of the gases Po is inclined at an angle a on the front of the missile to provide the two components of acceleration force: the longitudinal force F _L making it possible to compensate for the force of terrestrial gravity and the normal force F _N used in combination with the roll angle of the missile body to vary the orientation of the speed vector V of the missile. The section of the combustion chamber and, consequently, the section of the propellant blocks, can be toroidal in shape to allow free passage around the longitudinal axis X of the missile, in particular for arranging the coupling shaft 21 of the front and rear sections of the missile.

où F est la force de poussée nécessaire, Td la durée de trajet maximale du missile sur la portion terminale de sa trajectoire et I_s l'impulsion spécifique du propergol utilisé.The total mass mp of propellant must satisfy the following relationship:

where F is the necessary thrust force, Td the maximum travel time of the missile on the terminal portion of its trajectory and I _s the specific impulse of the propellant used.

The military charge can advantageously be of the so-called "hollow charge" type which produces a jet capable of puncturing the vehicle's protective armor. To ensure free passage of the jet along the longitudinal axis of the missile, the coupling shaft 21 of the front and rear sections of the missile comprises a recess in its axial portion; moreover, a free passage can also be arranged in the central part of the compartment 25 bringing together the electronic circuits associated with the sensor E.0 23 and with the drive member 24.

The braking parachute 35 of the missile can be a parachute similar to those used in the technique of braked projectiles such as aviation bombs. With this parachute are associated release and release devices not shown. The duration of action of the parachute is a function of the mass Mo of the missile and of the ratio of the cruising speed to the predetermined speed V over the terminal portion of the trajectory of the missile.

The guided missile which has just been described in detail may be a medium caliber missile of the order of 100 mm and an elongation factor of approximately 6 to 7 for a weight of 10 to 15 kgs. However, it can be indicated that all of its values can be modified within wide limits depending in particular on the destructive power of the military charge carried.

The guided missile, in itself, as just described, can constitute a sub-projectile of a projectile of larger dimensions whose main function is to ensure the carrying of this or a grouping of such sub-projectiles on the cruise portion to the end position of the firing trajectory.

Reference is now made to FIG. 14 which illustrates the transient portion between the cruising portion and the terminal portion of the firing trajectory. The carrier projectile 50 transports sub-projectiles or guided missiles 51, 52 and 53 located in a section 54. From the start of the transition portion of the trajectory, the guided missiles are ejected and dispersed with an important initial speed substantially equal to that of the carrier projectile and are at a predetermined altitude above the ground. In order to reduce their initial speed of movement to reach the speed V suitable for carrying out the acquisition and interception of targets, the braking parachute 35 of the missile is released for a determined period, after which the mechanical link between the missile and the parachute is broken to ensure the release of it. The stabilizing stabilizer 31 is deployed and the front section of the missile is put into autorotation. Consequently, the gas generator, to produce the transverse thrust force F _N is activated and the phase of search for a potential target situated on the ground can begin. It results from the ejection force imparted by the carrier vehicle 50 at the time of its separation from the sub-projectiles 51 to 53, a certain distance of dispersion R _D at the moment when the operation of search for the targets by the sensor begins. of the sub-projectile.

Figure 15 is a partial exploded view of section 54 of the carrier projectile 50 which shows an example of installation of a group of three guided missiles 51, 52 and 53. These missiles are regularly distributed around the longitudinal axis of the projectile carrier, moreover, an identical grouping of missiles can be installed in tandem, if necessary.

Figure 16 is a cross section of the carrier projectile 50 which shows the relative arrangement of the guided missiles 51, 52 and 53 inside the housing section 54. The guided missiles are supported on elements 55 actuated by a mechanism ejection 56 whose complementary function is to communicate a certain amount of movement to the missiles during their ejection, in order to ensure a predetermined relative dispersion. The ejection mechanism 56 can be of a known mechanical type actuated by hydraulic, pneumatic or possibly electrical means. In order to minimize the cross-section of the carrier projectile, the missiles can be provided with a tail unit in the form of four deployable fins 32, in order to allow a certain material embedding thereof.

The Table is a summary table of the progress of the main operations carried out by the missile during its firing trajectory.

The guided missile according to the invention is not limited in its characteristics and its applications to the embodiment described. In particular, the sensor can be of the passive or semi-active type and operate in the optical or radar bands of the electromagnetic spectrum, the relative arrangement of the elements such as the driving member 24 and the military charge 33 can be modified.

The invention is not limited to its application to an autonomous missile, but also applies to a missile carried by conventional vehicles or aircraft.

Claims (12)

1. A method for the guidance, during the last part of its trajectory, of a missile (10) having a first section (20) coupled by a central shaft (21) with a second section (30), the first section (20) comprising an engine means (24);

- the said engine means (24) and the central shaft (21) permitting relative rotation between two sections of the missile around a common axis, said axis being the longitudinal axis (X) of the body of the missile,

- the first section (20) having means (26) in order to create a transverse thrust force (F_N) normal to the direction of the speed (V) of displacement of the missile,

- the first section of the missile being provided with a sensor (23c) responsive to the energy radiated by a potential target,
said method comprising, for seeking the target, the following sequences:

- immobilizing the beam (14) of reception of the sensor on the longitudinal axis (X) of the missile; .

- imparting on the body of the missile a rotation with the given angular velocity of roll (0) around the longitudinal axis (X) of the missile;

- creating the said transverse thrust force (F_N) and maintaining the same force during the rest of the terminal phase in order to impart a helical motion on the body of the missile so that it spirally sweeps the said beam (14),

- detecting the image of any target sensed by the reception beam of the sensor (23c);
said method comprising, for pursuing the target after detection, the following stages:

- liberating the reception beam (14) of the sensor (23c);

- maintaining the axis (15) of this beam directed on the target which has been detected, the said axis of the beam forming a line (15) of aiming between the missile and the target;

- measuring the velocity of rotation (η) of the aiming line,

- elaborating an order of piloting (Vd proportional to the measured magnitude of the velocity of the said rotation (₁₁) of the aiming line,

- applying this order of piloting to the motor means (24) in such a manner that it causes the first section (20) to turn in relation to the second section (30) of the missile, with the purpose of modifying the angle of roll (0) of the first section, in order to cancel the rotation (₁₁) of the aiming line using the transverse force in such a manner that the direction of the trajectory of the missile towards the target be corrected.

2. The method of guiding as claimed in claim 1, characterized in that the velocity of displacement of the missile along its trajectory is established to be at a predetermined value (V) at the time when the latter commences the terminal part of its trajectory.

3. The method of guiding as claimed in claim 2, characterized in that during the terminal portion of its trajectory the missile performs a motion which is essentially vertical and that the velocity (V) of displacement of the missile along its trajectory and at the terminal portion of its trajectory is maintained essentially constant by creating a longitudinal thrust force (F_L) with a magnitude substantially equal to the resulting force of the field of terrestrial gravity (g) and the direction aligned with the longitudinal axis (X) of the missile.

4. The method of guiding as claimed in claim 3, characterized in that the angular velocity (0) of roll of the body of the missile is increased along the terminal portion of the trajectory of the missile during the seeking phase.

5. A guided missile, comprising a front section (20) coupled by a central shaft (21) with a rear section (30), the front section comprising an engine means (24), the said engine means (24) and the central shaft (21) permitting relative rotation of the two sections of the missile,

-the front section (20) furthermore comprising a section (23c) responsive to the energy radiated by a potential target and means to furnish a transverse thrust force (F_N),

- the engine means (24) comprising a control input connected with a generator of piloting orders,

- the missile comprising at its base a stabilizing tail (31) in the form of fins,
characterized in that:

- the two sections (20 and 30) of the missile are in a state of relative rotation about a common axis, said axis being the longitudinal axis (X) of the body of the missile,

- the sensor is provided on the one hand with a latching system to immobilize the reception beam (14) of the sensor (23c) along the longitudinal axis (X) of the missile during the phase of seeking, and on the other hand means to unlatch the sensor (23c) on detecting the target and in order keep the reception beam axis (14) directed on the detected target, the said beam axis forming a missile/target aiming line, an order of piloting being created in response to the velocity of rotation of the aiming line,

- the engine means (24) comprises a first member (24c) integral with the structure of the front section and a second member (24b) physically coupled with the rear section,

- the said control input is connected with the piloting generator by the intermediary of an amplifier,

- the engine means (24) modifies, in accordance with the piloting order, the roll angle (Φ) of the front section (20) during the pursuit phase,

- the means for providing the transverse thrust are constituted by a gas generator (26) supplying a lateral nozzle (27),

- the stabilizing fins are able to be deployed.

6. The missile as claimed in claim 5, characterized in that the second member (24b) of the engine means is mechanically coupled with the rear section (30) of the missile by the central coupling shaft (21).

7. The missile as claimed in claim 6, characterized in that the engine means (24) is an electrical clutched motor.

8. The missile as claimed in claim 7, characterized in that the rear section (30) of the missile comprises a military charge of the "hollow charge type" and in that the rear section (20) comprises an axial cavity (21 a).

9. The missile as claimed in any one of the claims 5 through 8, characterized in that the rear section (30) of the missile comprises a compartment (34) for the stowing of a parachute (35).

10. The missile as claimed in any one of the claims 5 through 9, characterized in that the stabilizing tail (31) is constituted by a set of fins (32) able to be folded back onto the body of the missile.

11. The missile as claimed in any one of the claims 5 through 10, characterized in that it constitutes a subprojectile of a carrying projectile.

Images