EP0482970A1 - Device for imparting a deviated trajectory to an airborne missile - Google Patents

Abstract

The method according to the invention consists in ensuring that the missile is provided with lift wings (4) and that when launched from an aircraft in flight, the lift wings (4) are inclined in a transverse direction in such a way that the load factor (f) has a horizontal component (fy) resulting in a lateral displacement of the missile. In the case of a self-propelled missile which has to fly substantially horizontally at a speed (V), the lift wings (4) are then designed such that for the speed (V) the vertical component (fz) of the load factor (f) cancels the effect of gravity. Applications to the launching of decoy rockets and to off-aim firing. <IMAGE>

Description

The present invention relates to the launching of a projectile from an aerial vehicle, in the direction of its trajectory. It relates more particularly to a method and a device for laterally bending the trajectory of the projectile relative to the direction of fire.

The invention applies in particular to the launching of a self-propelled projectile, such as a decoy rocket, from an airplane. The mission of such a rocket is to protect, by means of an on-board decoy, an airplane flying at very low altitude during the final attack phase of an objective defended by solair systems. Of course, the confrontation with means of detection and processing of information capable of distinguishing between different signatures requires that the rocket, to be credible, evolves within limits of which the aircraft is capable. This implies, among other things, that the rocket is launched along the trajectory of the aircraft, and that after launch, its relative speed with respect to it remains low, in practice of the order of 20 m / s. Under these conditions, the gap between the rocket and the plane does not grow quickly, and if the latter follows a straight path as is generally the case, this gap remains insufficient to avoid the risk of interception by incoming missiles. front.

The solution to this problem is to imprint a laterally inflected trajectory on the decoy rocket to attract the threat at a sufficient distance from the aircraft to be protected. For this purpose, we are currently using trajectory programming techniques which have the disadvantage of requiring expensive equipment, based on sensors and means for processing complex information.

The subject of the invention is another method for printing a projectile laterally deflected trajectory in a controlled manner after firing, as well as means for its implementation, which have the advantage of being simple, and therefore inexpensive, and to be easily applicable to the launching of a decoy rocket.

The method according to the invention is characterized in that it consists in providing that the projectile is provided with a lift wing, and that when launched from a vehicle in flight, said wing lift of the projectile is inclined transversely, so that the load factor has a horizontal component causing a lateral displacement of the projectile.

• une voilure de portance sur le projectile;
• une rampe de lancement sur l'engin porteur, dirigée horizontalement vers l'avant,
• des moyens pour retenir, le long de ladite rampe, le projectile dans une position angulaire déterminée autour de son axe longitudinal, correspondant à une inclinaison transversale donnée de la voilure de portance; et
• un moyen de propulsion pour assurer le lancement dudit projectile.

A device for implementing this method is characterized in that it comprises:

• an airfoil on the projectile;
• a launching ramp on the carrier, directed horizontally forward,
• means for retaining, along said ramp, the projectile in a determined angular position around its longitudinal axis, corresponding to a given transverse inclination of the lift wing; and
• a propulsion means for launching said projectile.

In the case of a self-propelled projectile such as a decoy rocket intended to perform a substantially horizontal flight, the airfoil is then calculated so that for the cruising speed of the projectile, the vertical component of the load factor cancels gravity.

• la Fig. 1 est une vue schématique en plan d'une roquette leurre faisant partie d'un système conforme à l'invention,
• les Figs. 2 et 3 sont respectivement une vue en coupe transversale et une vue partielle en coupe longitudinale de la roquette de la Fig. 1, illustrant schématiquement une forme de réalisation préférée de sa voilure de portance,
• la Fig. 4 est une demi-vue en coupe longitudinale de l'arrière de la roquette, illustrant une forme de réalisation préférée de l'empennage de queue,
• la Fig. 5 est un diagramme représentant les composantes verticale et transversale du facteur de charge sur la voilure de portance lorsque celle-ci est inclinée transversalement à 45°,
• la Fig. 6 est un diagramme semblable à celui de la Fig. 5, pour le cas où la voilure de portance est inclinée transversalement à 60°,
• la Fig. 7 est un graphique sur lequel apparaissent les trajectoires de l'avion et de la roquette, suite au lancement de cette dernière,
• la Fig. 8 est une demi-vue en coupe axiale représentant schématiquement un dispositif de détection de roulis,
• la Fig. 9 est une demi-vue en coupe latérale selon la ligne IX-IX de la Fig. 8, et
• les Figs. 10 et 11 sont des vues schématiques illustrant sous deux formes de réalisation différentes des moyens de correction de roulis complémentaires d'un dispositif de détection de roulis tel que celui des Figs. 8 et 9.

The invention will be better understood with the aid of the explanations which follow and of the appended drawings, in which:

• Fig. 1 is a schematic plan view of a decoy rocket forming part of a system according to the invention,
• Figs. 2 and 3 are respectively a cross-sectional view and a partial view in longitudinal section of the rocket of FIG. 1, schematically illustrating a preferred embodiment of its lift wing,
• Fig. 4 is a half view in longitudinal section of the rear of the rocket, illustrating a preferred embodiment of the tail fin,
• Fig. 5 is a diagram representing the vertical and transverse components of the load factor on the lift wing when the latter is inclined transversely at 45 °,
• Fig. 6 is a diagram similar to that of FIG. 5, in the case where the lift wing is inclined transversely at 60 °,
• Fig. 7 is a graph on which the trajectories of the aircraft and the rocket appear, following the launch of the latter,
• Fig. 8 is a half-view in axial section schematically showing a roll detection device,
• Fig. 9 is a half-view in lateral section along the line IX-IX in FIG. 8, and
• Figs. 10 and 11 are schematic views illustrating in two different embodiments means for correcting roll complementary to a roll detection device such as that of FIGS. 8 and 9.

The projectile shown in the drawings is a decoy rocket intended to be launched from an airplane, substantially in the direction of its speed vector when it is in horizontal flight.

The rocket comprises a front section 1 containing the lure (which can be of any type, for example electromagnetic, pyrotechnic or thermal), an intermediate section 2 at the level of which is the wing structure made up of two wings 4, and a rear section or propulsive part 3. On the section 3, a tail fin is installed formed by a plurality of fins 5.

Preferably, the fins 5, as well as the wings 4 forming the lift wing, are deployable, which makes it possible to have all the elements of the rocket included in its caliber and to reduce the aerodynamic constraints in flight before launching.

Figs. 2 and 3 schematically represent a simple deployable wing system, which ensures an opening conjugation of the two wings 4. These are articulated so as to be able to be fully retracted inside the section 2, in the position shown in dashed lines in FIG. 3, and they comprise pinions 6 meshed with the two sides forming a rack of a flat rod 7. The longitudinal displacement of the rod 7 necessary for the combined opening of the wings 4 is ensured, for example, by spring means and lock released when launching the rocket.

In the embodiment shown, the propulsion section 3 comprises two separate parts: either a take-off stage with the load 8 and the peripheral nozzles 9, which stage must provide the acceleration necessary for launching, and a second cruising stage with the load 10 and the central nozzle 11, which has for function of maintaining the rocket at a predetermined speed substantially constant, after launch.

The rocket takes place on a plane launching ramp, directed forward and oriented so as to be parallel to the speed vector when the plane is in horizontal flight. The ramp may be a rail, or else a tube when the lift wing and the tail of the rocket are deployable, and included in its caliber in the folded position. Note that in this second case, the tube may be a cannon, the take-off thruster 8.9 then no longer having any reason for being.

Means are provided to ensure the positioning of the rocket on the ramp in a determined position of rotation about its longitudinal axis, and also to maintain this position throughout the ramp at the time of launch. In the case of a launching tube, these means comprise, for example, a finger secured to the rocket, which can be engaged in a guide groove made in the inner wall of the tube.

According to the invention, the rotational position of the rocket on the launching ramp defines a determined angle α of transverse inclination of its lift wing (wings of the aircraft in horizontal position). On the other hand, said lift wing is calculated so that under these conditions of transverse inclination, and for a speed V slightly higher than the recommended speed of the aircraft at the time of the firing, the load factor resulting from the lift and of the thrust component of the cruise thruster is oriented perpendicular to the axis of the rocket and has a vertical component which cancels gravity. The thrust generated by the cruising thruster 10, 11 is approximately equal to the drag of the rocket, so that this speed V is kept constant, or slightly increasing.

Fig. 5 represents the load factor f on a graph (Oy, 0z) for a transverse tilt angle α equal to 45 °. The projections f _y and f _z are then equal to g, which implies for f the value g√2.

Under these conditions, and insofar as the angle α remains constant, the rocket follows after launch a circular horizontal trajectory of radius R = V² / g. This trajectory is represented on the graph (Ox, Oy) of FIG. 7, C being the center of the trajectory, the origin 0 representing the position of the airplane at the time of launch and Ox the direction of its trajectory.

d'où :
   ϑ = VT/R = gT/V, et
   ML = R sinϑ = V²/gsinϑ
   OM = V²/g(1 - cosϑ) (M étant la projection de L sur 0y)At the end of a time T, the rocket arrives at L while the plane which has followed a rectilinear trajectory arrives at A. We then have:

from where :
ϑ = VT / R = gT / V, and
ML = R sinϑ = V² / gsinϑ
OM = V² / g (1 - cosϑ) (M being the projection of L on 0y)

Taking the following values as an example:
Va (airplane speed) = 300 m / s
V (rocket speed) = 320 m / s
T = 10 s,
we deduce :
ϑ = 17.5 °
OA = V _a T = 3000m
ML = 3,180 m
OM = 487 m

The value of the transverse angle of inclination α can be chosen within a range limited below by the need to obtain a sufficient lateral deviation, and above by the credibility of evolution of the rocket and by the excessive increase in the factor of charge f. As an indication, Fig.6 illustrates this for a transverse tilt angle α of 60 °: f _z = g then implies f = 2g and f _y = 1.7g. In practice, as the previous calculation shows, the value of 45 °, in addition to its geometric simplicity, represents an interesting compromise.

In reality, the transverse inclination of the lift wing does not remain constant, since the rocket is generally subject to a roll drive which should be reduced as much as possible, or corrected according to the applications.

In cases where the propulsion time and, consequently, the range of the rocket are short, the random variation of the angle α of transverse inclination can lead to an acceptable deviation from the estimated trajectory. There is therefore no need to introduce a correction.

d'où z = 1/2gdαt²The following simplified calculation makes it possible to evaluate an upper limit of the trajectory alteration: if the angle α deviates from the start of the nominal value 45 ° (π / 4) from an algebraic increase dα assumed to be constant, l vertical acceleration applied to the rocket becomes:

hence z = 1 / 2gdαt²

For dα = 1 ° (0.017 radian) and t = 5 s, we find:
z = 2.14 m

We see that for a variation of α not exceeding a few degrees, the resulting variation in altitude remains acceptable. The criterion is the risk that the rocket hits the ground too early because, in practice, this type of projectile is launched at very low altitude, of the order of 30 m.

The lateral alteration (according to Oy in Fig. 7) is less important since we are not looking for a precise trajectory, but only an evasive one which credibly simulates that of the firing plane. This alteration would result in a maximum variation of approximately 2 m for dα = 1 °, t = 5 s and V = 320 m / s.

This example makes it possible to estimate the order of magnitude of the expected alterations. In reality, the hypothesis according to which dα is constant is penalizing because α varies according to a more or less alternative law insofar as the asymmetries of shape have been reduced by the quality of manufacture and where a free tail in rotation is provided around the rocket, because it is he who is responsible for most of the roll training.

Fig. 4 illustrates such an arrangement, in which the propulsion section 3 comprises a part 3a joined to the intermediate section 2 and housing the propellant cruise 10, 11, and a part 3b in the form of a cylindrical block mounted to rotate freely, by means of bearings 12, around the part 3a. Part 3b comprises the take-off thruster 8.9, as well as means for installing the tail unit, that is to say in the example shown of the legs 13 on which the fins 5 are articulated. As a variant, only the tail unit is mounted on bearings around section 3.

If the deviations of the angle α and their consequence on the trajectory are unacceptable, which may be the case when the propulsion time is relatively long, it becomes necessary to provide means for correcting roll. A simple and therefore inexpensive example of such means is shown in Figs. 8 to 11. A flywheel 14 is mounted to rotate freely around the axis of the rocket, so as not to be driven by the roll deviations of the latter. The friction can be considered negligible insofar as the stresses are alternative. In practice, this tendency is favored, to the detriment of stresses repeating in the same direction, by providing a free tail in rotation as described above.

The steering wheel 14, while remaining stationary in rotation, serves as a reference for assessing the deviations in roll of the rocket. Figs. 8 and 9 illustrate means by which signals are produced when these deviations reach a determined value dα in one or the other direction. On the flywheel 14, these means consist of a through channel 15, located in a radial plane r, at a distance d from the axis.

Attached to the rocket body, two sets each include a light source 16 on one side of the steering wheel 14, and a photoelectric cell 17 on the other side. As seen in Fig. 9, these two assemblies are also located at a distance d from the axis, and are angularly separated by the value dα with respect to the longitudinal plane of symmetry of the lift wing formed by the wings 4. In addition, means are necessary for bring the radial plane r of the flywheel 14 to coincide with this plane of symmetry, between the two assemblies, FIG. 9. The steering wheel 14 is retained in this position by a lock released at launch.

The angular spacing dα of the assemblies with respect to said plane of symmetry corresponds to the maximum roll angle deviation that is tolerated. When this maximum deviation is reached on one side or the other, as illustrated in FIG. 8, the corresponding photoelectric cell receives light from its light source and produces a signal that can be used to control a roll correction device, two embodiments of which are shown by way of example in FIGS. 10 and 11.

In Fig. 10, a reserve of pressurized gas 18 is connected by conduits 19 to peripheral nozzles 20. They are preferably associated in pairs, each of which is capable of producing counter-roll pulses in one direction, following the selective and time-controlled opening of solenoid valves 21.

The device of FIG. 11 is made up of small cannons 22, also associated in pairs, preferably. By launching projectiles, these guns produce recoil forces which can be used as counter-roll pulses.

When the processes described above become insufficient because we seek higher precision, or that we work over a longer time, or for another reason, we must call on components more sophisticated such as an inertial unit, but which are obviously much more expensive.

The field of use of the invention is not limited to the launching of decoy rockets which has just been described. Among other interesting applications, it is possible to point the shot in aiming since a high precision is not required: either for example for a zone treatment by rocket bursts, the advantage being that the plane can operate at low altitude without having to pass above the objective. In this application, in the event of required trajectory horizontality, the ranges of possible values for the speed V and the angle α of transverse inclination of the lift wing are obviously much greater than for the case of a decoy rocket. .

Claims (12)

Projectile having a longitudinal axis of the type comprising a lift car (4) a tail fin (5) and intended to be released from a carrier air vehicle, the use of the projectile comprising a flight phase carried on the carrier vehicle and an autonomous flight phase from the drop, characterized in that it comprises: means for holding the projectile in a roll position inclined at an angle α relative to the horizontal during the flight phase carried, these means being intended to cooperate for this purpose with means complementary to the missile carrier ; - means for holding the projectile in a roll position inclined relative to the horizontal during the autonomous flight phase; - a cruise thruster (10, 11). Projectile according to claim 1, characterized in that the means for holding the projectile in an inclined roll position during the autonomous flight phase comprises means for detecting a predetermined deviation dα between the real value of the angle d 'inclination of the airfoil and the expected value α and means for correcting this deviation controlled by the detection means. Projectile according to claim 2, characterized in that the deviation detection means are constituted by a flywheel (14) comprising two faces and having as axis the axis of the projectile, the flywheel comprising a through hole (15) located in a radial plane r at a distance d from the axis. - Two light sources (16) integral with the projectile illuminating a first face of the steering wheel, the two sources being at a distance d from the axis of the projectile and in radial planes forming an angle between them equal to twice the predetermined value dα. - Two photoelectric cells (17) symmetrical of the light sources with respect to the steering wheel. - Retractable means for holding the steering wheel, in a position where the radial plane r of the through hole (15) makes an angle dα with each of the radial planes passing through the photoelectric cells (17) and the light sources (16). Projectile according to claim 2, characterized in that the roll correction means consist of a reserve of pressurized gas (18) connected by conduits (19) to peripheral nozzles (20) the opening and closing of the reserve being controlled by at least one solenoid valve (21). Projectile according to claim 2, characterized in that the roll correction means consist of small cannons (22) capable of producing, in the event of a pull, recoil forces usable as a counter-roll impulse. Projectile according to one of claims 1 to 5, characterized in that it further comprises a take-off propellant (8, 9). Projectile according to one of claims 1 to 5, characterized in that the tail of the tail (5) is mounted free to rotate about its longitudinal axis. Projectile according to Claim 6, characterized in that the tail of the tail (5) and the take-off propellant (8, 9) form an assembly mounted free in rotation around the longitudinal axis of the projectile. Projectile according to one of claims 1 to 8, characterized in that the lift wing (4) and the tail fin (5) are deployable. Projectile according to one of Claims 1 to 9, characterized in that the cruising thruster is intended to maintain the projectile at a predetermined speed V and in that at this speed V, the component (fz) of the load factor f applied to the projectile inclined by the angle α in roll cancels the weight of the projectile. Projectile according to one of claims 1 to 10, characterized in that it carries electromagnetic or infrared decoys. Projectile according to claim 11, characterized in that it is organized in three sections, a front section (1) containing the lures, an intermediate section (2) comprising the lift wing (4) and a rear section (3) comprising the propulsive part and tail fin.

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