EP0081421B1 - Verfahren zur Endphasenlenkung und dieses Verfahren verwendender Lenkflugkörper - Google Patents

Verfahren zur Endphasenlenkung und dieses Verfahren verwendender Lenkflugkörper Download PDF

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Publication number
EP0081421B1
EP0081421B1 EP82402180A EP82402180A EP0081421B1 EP 0081421 B1 EP0081421 B1 EP 0081421B1 EP 82402180 A EP82402180 A EP 82402180A EP 82402180 A EP82402180 A EP 82402180A EP 0081421 B1 EP0081421 B1 EP 0081421B1
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EP
European Patent Office
Prior art keywords
missile
section
target
sensor
trajectory
Prior art date
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Expired
Application number
EP82402180A
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English (en)
French (fr)
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EP0081421A1 (de
Inventor
Pierre Metz
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Thomson Brandt Armements SA
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Thomson Brandt Armements SA
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Priority to AT82402180T priority Critical patent/ATE40467T1/de
Publication of EP0081421A1 publication Critical patent/EP0081421A1/de
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/60—Steering arrangements
    • F42B10/66—Steering by varying intensity or direction of thrust
    • F42B10/661—Steering by varying intensity or direction of thrust using several transversally acting rocket motors, each motor containing an individual propellant charge, e.g. solid charge
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/20—Direction control systems for self-propelled missiles based on continuous observation of target position
    • F41G7/22—Homing guidance systems
    • F41G7/222—Homing guidance systems for spin-stabilized missiles
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/20—Direction control systems for self-propelled missiles based on continuous observation of target position
    • F41G7/22—Homing guidance systems
    • F41G7/2233—Multimissile systems
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/20—Direction control systems for self-propelled missiles based on continuous observation of target position
    • F41G7/22—Homing guidance systems
    • F41G7/2253—Passive homing systems, i.e. comprising a receiver and do not requiring an active illumination of the target
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41G—WEAPON SIGHTS; AIMING
    • F41G7/00—Direction control systems for self-propelled missiles
    • F41G7/20—Direction control systems for self-propelled missiles based on continuous observation of target position
    • F41G7/22—Homing guidance systems
    • F41G7/2273—Homing guidance systems characterised by the type of waves
    • F41G7/2293—Homing guidance systems characterised by the type of waves using electromagnetic waves other than radio waves
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42C—AMMUNITION FUZES; ARMING OR SAFETY MEANS THEREFOR
    • F42C13/00—Proximity fuzes; Fuzes for remote detonation
    • F42C13/006—Proximity fuzes; Fuzes for remote detonation for non-guided, spinning, braked or gravity-driven weapons, e.g. parachute-braked sub-munitions

Definitions

  • 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.
  • 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.
  • the missile may be provided with a military charge capable of perforating the protective armor of armored vehicles.
  • the senor 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.
  • 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.
  • 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.
  • 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.
  • the search head 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.
  • 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.
  • 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.
  • 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.
  • the proposed guidance method uses a target tracking sensor which measures the rotation of the missile / target line of sight.
  • 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.
  • the second member of the engine member is coupled to the rear section of the missile by a central shaft.
  • 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.
  • 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 ⁇ r around its longitudinal axis X carrying the displacement speed vector V of the projectile on its path.
  • the trajectory of the projectile is inclined by an angle ⁇ t and this projectile strikes the ground at a point 4 angularly offset by an angle ⁇ c of a potential target 6.
  • the projectile 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 ⁇ 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.
  • 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.
  • 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 ⁇ 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 ⁇ 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.
  • 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 ⁇ N normal to the speed vector V, the longitudinal force F L to which corresponds an acceleration ⁇ 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 ⁇ , 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 + ⁇ ) 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 ⁇ o remains high, 60 to 70 ° for example.
  • 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 ⁇ o equal to 90 °.
  • 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 ⁇ A s intercepted by the optical beam is an ellipse whose magnitudes of the axes ⁇ R s and ⁇ R ' s are given respectively by the following relationships:
  • a target c animated with a speed V c and distant by a value R e from point I has also been indicated.
  • the angular speed ⁇ 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.
  • the transit time T D of the optical beam on a target c is given by the following relation: where ⁇ 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:
  • the speed component V t is tangent to the circle with center 0 and radius r.
  • the value of the angle of inclination ⁇ of the speed vector V of the missile is obtained, with respect to the generator GI of the cylinder.
  • FIG. 8 is a simplified diagram representing a variant of the method of searching for a target on the ground.
  • 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 ⁇ R s of the successive scanning frames and the angle of inclination ⁇ 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 ⁇ A s considering that the values of the angles ⁇ and ⁇ have always small values.
  • the trajectory S of the center of gravity G of the missile is inscribed on the surface of a cone.
  • 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 ⁇ 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 ⁇ 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.
  • the equation for the missile piloting law is of the form: which corresponds to a proportional gain navigation law A comprising a bias ⁇ o , ⁇ ⁇ 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 ⁇ o given by the following relation:
  • the piloting input signal is proportional to the quantity ⁇ and the response is the quantity A0 of the orientation of the thrust force F N relative to the direction of the rotation vector ⁇ such that since the terms ⁇ 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 ⁇ .
  • 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 ⁇ y and ⁇ 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
  • the crossed component of the acceleration y T ⁇ 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 ⁇ 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 ⁇ 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.
  • 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.
  • 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 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 ⁇ , 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.
  • 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.
  • 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.
  • 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 receives an electrical signal resulting from the comparison of the angular velocity ⁇ 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • FIG. 14 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.
  • 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.
  • 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.

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  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Electromagnetism (AREA)
  • Fluid Mechanics (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)

Claims (12)

1. Verfahren zur Lenkung eines Flugkörpers (10) während des Endbereichs seiner Flugbahn, wobei dieser Flugkörper einen ersten Abschnitt (20) hat, der über einen zentralen Schaft (21) an einen zweiten Abschnitt (30) angekoppelt ist, und der erste Abschnitt (20) ein Antriebsorgan (24) umfaßt:
- wobei dieses Antriebsorgan (24) und dieser zentrale Schaft (21) zwischen den beiden Abschnitten des Flugkörpers eine Relativdrehung um eine selbe Achse ermöglichen und diese Achse die Längsachse (X) des Rumpfes des Flugkörpers ist,
- wobei der erste Abschnitt (20) Mittel (26) besitzt, um eine quergerichtete Schubkraft (FN) zu erzeugen, welche zu der Richtung der Fortbewegungsgeschwindigkeit (V) des Flugkörpers senkrecht ist,
- wobei der erste Abschnitt des Flugkörpers mit einem Sensor (23c) versehen ist, der für die von einem potentiellen Ziel ausgestrahlte Energie empfindlich ist,
wobei das Verfahren die folgenden Sequenzen für die Zielsuchphase umfaßt:
- Stillsetzen des Empfangsbündels (14) des Sensors auf die Längsachse (X) des Flugkörpers;
- dem Rumpf des Flugkörpers wird eine Drehung mit bestimmter Rollwinkelgeschwindigkeit (0) um die Längsachse (X) des Flugkörpers aufgegeben;
- Erzeugen der genannten quergerichteten Schubkraft (FN) sowie Aufrechterhaltung derselben Kraft während der übrigen Endphase, um dem Rumpf des Flugkörpers eine Schraubenbewegung aufzugeben, so daß er eine spiralförmige Verschwenkung des genannten Bündels (14) ausführt, und
- Erfassen des Bildes eines durch das Empfangsbündel des Sensors (23c) eventuell aufgefangenen Ziels;
wobei das Verfahren die folgenden Schritte für die Zielverfolgung nach der Erfassung umfaßt:
- Freigabe des Empfangsbündels (14) des Sensors (23c);
- Halten der Achse (15) dieses Bündels auf dem erfaßten Ziel, wobei die genannte Bündelachse eine Flugkörper/Ziel-Visierlinie (15) bildet;
- Messen der Drehgeschwindigkeit (η) der Visierlinie;
- Erzeugen eines Lenkbefehls (VE), der zu der gemessenen Größe der Geschwindigkeit der gennanten Drehung (η) der Visierlinie proportional ist;
- diesen Lenkbefehl an das Antriebsorgan (24) anlegen, so daß es den ersten Abschnitt (20) in Bezug auf den zweiten Abschnitt (30) des Flugkörpers in Drehung versetzt, mit dem Ziel, den Rollwinkel (0) des ersten Abschnitts zu verändern, im die Drehung (η) der Visierlinie durch Einwirkung der quergerichteten Kraft zu annullieren, so daß die Richtung der Flugbahn des Flugkörpers gegen das Ziel korrigiert wird.
2. Verfahren zur Lenkung nach Anspruch 1, dadurch gekennzeichnet, daß die Fortbewegungsgeschwindigkeit des Flugkörpers entlang seiner Flugbahn auf einem bestimmten Wert (V) zu dem Zeitpunkt festgelegt wird, wo dieser den Endbereich seiner Flugbahn anfliegt.
3. Verfahren zur Lenkung nach Anspruch 2, dadurch gekennzeichnet, daß der Flugkörper eine im wesentlichen senkrechte Bewegung während des Endbereichs seiner Flugbahn hat und daß die Fortbewegungsgeschwindigkeit (V) des Flugkörpers entlang seiner Flugbahn und auf dem Endbereich derselben im wesentlichen dadurch konstant gehalten wird, daß eine Längsschubkraft (FL) erzeugt wird, deren Größe der aus dem Erdschwerekraftfeld (g) resultierenden Kraft entspricht und deren Richtung auf die Längsachse (X) des Flugkörpers ausgerichtet ist.
4. Verfahren zur Lenkung nach Anspruch 3, dadurch gekennzeichnet, daß die Rollwinkelgeschwindigkeit (0) des Rumpfes des Flugkörpers während der Suchphase entlang des Endbereichs der Flugbahn des Flugkörpers erhöht wird.
5. Gelenkter Flugkörper mit einem Vorderabschnitt (20), der über einen zentralen Schaft (21) an einen Hinterabschnitt (30) angekoppelt ist, wobei der Vorderabschnitt ein Antriebsorgan (24) umfaßt, das genannte Antriebsorgan (24) sowie der zentrale Schaft (21) eine Relativdrehung zwischen den beiden Abschnitten des Flugkörpers ermöglichen,
- wobei der Vorderabschnitt (20) ferner einen Sensor (23c), der für die von einem potentiellen Ziel ausgestrahlte Energie empfindlich ist, sowie Mittel zum Erzeugen einer quergerichteten Schubkraft (FN), umfaßt,
- das Antriebsorgan (24) einen Steuereingang, der mit einem Lenkbefehlgenerator verbunden ist, umfaßt,
- der Flugkörper an seiner Basis ein stabilisierendes, als Flügel ausgebildetes Leitwerk (31) umfaßt,
dadurch gekennzeichnet, daß:
- die beiden Abschnitte (20, 30) des Flugkörpers um eine selbe Achse in Relativdrehung stehen, wobei diese Achse die Längsachse (X) des Flugkörpers ist,
- der Sensor einerseits mit einer Verriegelungsvorrichtung versehen ist, um während der Suchphase das Empfangsbündel (14) des Sensors (23c) entlang der Längsachse (X) des Flugkorpers festzuhalten, und andererseits mit Mitteln versehen ist zur Entriegelung des Sensors (23c) bei der Erfassung des Zieles und zum Festhalten des Empfangsbündels (14) auf dem erfaßten Ziel, wobei die genannte Bündelachse eine Flugkörper/Ziel-Visierlinie bildet und ein Lenkbefehl als Antwort auf die Drehgeschwindigkeit der Visierlinie gebildet wird.
- das Antriebsorgan (24) ein erstes Element (24c) umfaßt, das mit der Struktur des Vorderabschnittes festverbunden ist, sowie ein zweites Element (24b) umfaßt, das an den Hinterabschnitt direkt angekoppelt ist,
- der genannte Befehlseingang über einen Verstärker mit dem Lenkbefehlgenerator verbunden ist,
- das Antriebsorgan (24) den Rollwinkel (0) des Vorderabschnittes (20) während der Verfolgungsphase und gemäß des Lenkbefehls verändert,
- die Mittel zum Erzeugen der quergerichteten Schubkraft aus einem Gasgenerator (26) bestehen, der ein seitliches Düsenrohr (27) speist,
- die Stabilisierungsflügel entfaltbar sind.
6. Flugkörper nach Anspruch 5, dadurch gekennzeichnet, daß das zweite Element (24b) des Antriebsorgans an den Hinterabschnitt (30) des Flugkörpers über den zentralen Schaft (21) zur Ankoppelung mechanisch gekoppelt ist.
7. Flugkörper nach Anspruch 6, dadurch gekennzeichnet, daß das Antriebsorgan (24) ein elektrischer Drehmoment-Motor ist.
8. Flugkörper nach Anspruch 7, dadurch gekennzeichnet, daß der Hinterabschnitt (30) des Flugkörpers eine militärische Ladung vom Typ mit "Hohlladung" umfaßt und daß der Hinterabschnitt (20) eine axiale Aussparung (21 a) umfaßt.
9. Flugkörper nach einem der Ansprüche 5 bis 8, dadurch gekennzeichnet, daß der Hinterabschnitt (30) des Flugkörpers in Abteil (34) zur Aufnahme eines Fallschirmes (35) umfaßt.
10. Flugkörper nach einem der Ansprüche 5 bis 9, dadurch gekennzeichnet, daß das Stabilisierungsleitwerk (31) aus einem Satz von Flügeln (32) besteht, welche gegen den Rumpf des Flugkörpers einziehbar sind.
11. Flugkörper nach einem der Ansprüche 5 bis 10, dadurch gekennzeichnet, daß er ein Subprojektil eines Trägerprojektils darstellt.
EP82402180A 1981-12-09 1982-11-30 Verfahren zur Endphasenlenkung und dieses Verfahren verwendender Lenkflugkörper Expired EP0081421B1 (de)

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FR8123025A FR2517818A1 (fr) 1981-12-09 1981-12-09 Methode de guidage terminal et missile guide operant selon cette methode

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US4568040A (en) 1986-02-04
FR2517818A1 (fr) 1983-06-10
IL67424A (en) 1989-03-31
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CA1209232A (en) 1986-08-05
ATE40467T1 (de) 1989-02-15
EP0081421A1 (de) 1983-06-15
FR2517818B1 (de) 1985-02-22
DE3279397D1 (en) 1989-03-02
JPS58127100A (ja) 1983-07-28

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