EP0737297B1 - Missile launching and steering system - Google Patents

Missile launching and steering system Download PDF

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Publication number
EP0737297B1
EP0737297B1 EP95936617A EP95936617A EP0737297B1 EP 0737297 B1 EP0737297 B1 EP 0737297B1 EP 95936617 A EP95936617 A EP 95936617A EP 95936617 A EP95936617 A EP 95936617A EP 0737297 B1 EP0737297 B1 EP 0737297B1
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EP
European Patent Office
Prior art keywords
annular body
nozzles
missile
launching
gas
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
EP95936617A
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German (de)
French (fr)
Other versions
EP0737297A1 (en
Inventor
Ivan Ivanovitch Arkhangelsky
Eugène Gueorguevitch BOLOTOV
Vladimir Sergueevitch Thomson-CSF SCPI PHILIPPOV
Vladimir Yakovlevitch Thomson-CSF SCPI MIZROKHI
Vladimir Grigorievitch Svetlov
Gregory Andreevitch Stanevsky
Serge Grigorievitch Khitenkov
Victor Leonidovitch Gaidoukevitch
Eugène Afanassievitch CHMIKOV
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Thales SA
Le Bureau de Constuctions Mecaniques "FAKEL"
Original Assignee
Thomson CSF SA
Le Bureau de Constuctions Mecaniques "FAKEL"
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Filing date
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Priority claimed from RU95110350A external-priority patent/RU2082946C1/en
Application filed by Thomson CSF SA, Le Bureau de Constuctions Mecaniques "FAKEL" filed Critical Thomson CSF SA
Publication of EP0737297A1 publication Critical patent/EP0737297A1/en
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Publication of EP0737297B1 publication Critical patent/EP0737297B1/en
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    • 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
    • F42B10/663—Steering by varying intensity or direction of thrust using a plurality of transversally acting auxiliary nozzles, which are opened or closed by valves
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41F—APPARATUS FOR LAUNCHING PROJECTILES OR MISSILES FROM BARRELS, e.g. CANNONS; LAUNCHERS FOR ROCKETS OR TORPEDOES; HARPOON GUNS
    • F41F3/00—Rocket or torpedo launchers
    • F41F3/04—Rocket or torpedo launchers for rockets
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41F—APPARATUS FOR LAUNCHING PROJECTILES OR MISSILES FROM BARRELS, e.g. CANNONS; LAUNCHERS FOR ROCKETS OR TORPEDOES; HARPOON GUNS
    • F41F3/00—Rocket or torpedo launchers
    • F41F3/04—Rocket or torpedo launchers for rockets
    • F41F3/077—Doors or covers for launching tubes
    • 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

Definitions

  • the present invention relates to launching systems for flying machines, and in particular to missile launching and orientation systems. It can find its use for missiles of small or large dimensions, of the "ground-air” or “air-air” or “ground-ground” type.
  • Any system for launching and orienting flying machines comprises electronic control and supply means, as well as means necessary for the implementation of launching and orientation (mechanical, pyrotechnic means, etc.) under the control of said electronic means.
  • a missile launch and orientation system is known from US Pat. Nos. 3,286,956 and US-A-2,995,319 (base for the preamble of claim 1), which comprises launching means, aerodynamic control surfaces with their driving devices, as well as orientation means essentially comprising a gas generator and nozzles connected thereto.
  • the hot gases arrive from the gas generator which is located in the body of the missile, through the axes of rotation of the control surfaces, towards nozzles located in the rear part of the control surfaces and forming reactive jets directed parallel to the control surfaces.
  • the gas generator which is located in the body of the missile, through the axes of rotation of the control surfaces, towards nozzles located in the rear part of the control surfaces and forming reactive jets directed parallel to the control surfaces.
  • the launch and orientation system under consideration does not fully use the energy of the reactive jet, a jet which is parallel to the plane of the control surfaces, which decreases the angular speed of the missile when it changes direction towards the target. .
  • a system for launching and orienting missiles is known (International Patent WO 94/10527) which comprises launching means, aerodynamic control surfaces with their drive means, and orientation means comprising gas generators as well as nozzles connected to them.
  • this known system comprises a gas generator which is connected via gas pipes to pairs of nozzles; each pair is formed by two identical nozzles, oriented in opposite directions, the inlet ports of which face the outlet of their common gas pipe, and the diameters of which are identical to that of the outlet of the gas line.
  • This known system ensures the possibility of a rapid turn of the missile towards the target, thanks to the reactive jet ejected from each pair of nozzles, and perpendicular to the plane of the control surfaces.
  • orientation means form a block in common with the drive means for the control surfaces, which is difficult to integrate into the design of weak missiles. dimensions without degrading their aerodynamic properties. In addition, this excludes the possibility of dropping, after the missile has turned in the required direction, the inert mass represented by the orientation means.
  • This system could also be used for the above-mentioned modernization of inclined-launch missiles.
  • control system described in the article by Roger P. Berry, "Development of an orientation control system of the advanced kinetic energy missile” (ADKEM), AIAA-92-2763, also includes launching means, aerodynamic control surfaces with drive, as well as orientation means intended to be installed in the rear part of the missile, and the production of which is based on gas generators connected to nozzles.
  • the system described in this article can be adapted to missiles with inclined launch (to carry out the modernization mentioned above) without considerable modification of these missiles.
  • This system provides for the release of the inert mass of the orientation means after the performance of their function.
  • the complexity of the system, the large size orientation means which are provided exclusively for the use of highly toxic liquid fuels (hydrazine), make the implementation of this system very difficult.
  • orientation means are located on the path of the gases ejected by the nozzles of the cruise engines of the missile, it is necessary to provide for the release of the orientation means immediately after the turn towards the target. Furthermore, this release must be carried out immediately before the ignition of the cruise engines, that is to say above the launch pad, which complicates the execution of military actions and is also dangerous for the objective to defend.
  • None of the missile launch and orientation systems mentioned above can ensure the interception of a close target in the difficult conditions of a vertical departure, for example from the area located in a massif forest. This is first of all linked to the production of the means for launching these systems, means which do not allow a height of the order of 40 m to be quickly reached, necessary for perfectly performing the maneuvers of orientation towards the target and the ignition of the cruise engine.
  • the main problem to be solved by the present invention is the realization of a universal missile launch and orientation system, which it would be possible to combine both large and small missiles, allowing the launching of the inert mass of the orientation means far enough from the launch pad.
  • This system must be as inexpensive as possible and must be able to be used for all missiles with inclined departure, and must be able to provide omnidirectional defense.
  • the launching and orientation system of flying machines comprises launching means, aerodynamic control surfaces with their drive and orientation means, located in the rear part of the flying vehicle and comprising at least a gas generator and nozzles connected to it, and this system is characterized in that it comprises an annular body rigidly connected to the body of the flying object, the orientation means being located in the annular body , the internal surface of the annular body having a frusto-conical shape and being coated with a heat-insulating material, forming a nozzle section whose profile is in continuity with the profile of the nozzle of the cruising engine of the flying machine.
  • the annular body may include means ensuring its ejection by the flying object during the flight, which makes it possible to optimize the energy balance and to entirely release the inert mass which the orientation means represent after their use, at an instant. chosen, outside the launch pad area.
  • the nozzles of the orientation means are located in the same plane, perpendicular to the longitudinal axis of the nozzle section. This ensures optimal use of the energy of the reactive jets during the orientation of the flying object, and, consequently, allows the interception of the target near the launch pad.
  • the launch means are produced in the form of a launch container with front and rear covers, the internal volume of which has a cylindrical shape and is intended to receive the flying machine, the pressure generator being located at the bottom of the container, closed by a rear cover and by a protective shutter having a frustoconical lateral surface, the profile of which reproduces at least certain parts of the surface of the nozzle section of the annular body.
  • the rear part of the annular body has a peripheral valve, the outside diameter of which is equal to the inside diameter of the container.
  • the container has a support on which are weakened elements intended for fixing the annular body above the outlet openings of the pressure generator.
  • the protective shutter has a convex shape oriented towards the cruising engine. This embodiment of the shutter ensures, as described below, maximum reliability and efficiency of its operation in the launch system.
  • the launch container may include, in the fixing part of the annular body, an ejection orifice, the dimensions of which are chosen taking into account the gas flow rate passing through the clearance which is formed around the valve of the annular body.
  • the front cover of the container is made so as to be fragmented for a given pressure developing inside the container.
  • the launching and orientation system of the flying machine can be provided with rods fixed to the annular body.
  • the gas generator is also annular and connected to the nozzles of the orientation means by gas pipes formed in the annular body, the nozzles being all identical, grouped in pairs in the same plane.
  • the nozzles of each pair are oriented in opposite ways and mechanically connected to one end of the corresponding rod, which ensures the distribution of the gas jet between them from the common gas pipe of the annular body.
  • Each rod is connected by its other end to a corresponding control surface thus ensuring the possibility of a joint rotation. Consequently, the rotation of the aerodynamic control surfaces and of the orientation means is controlled by a single drive means.
  • the present invention provides two variants of the first embodiment of the launching and orientation system of the flying machine.
  • the control system is provided with annular sleeves made of heat-resistant material located near the outlet end of each corresponding gas pipe, these sleeves being able to move longitudinally.
  • Each rod is fixed to the annular body in its middle part by its axis of rotation.
  • Each pair of nozzles is produced in the form of bent pipes with frustoconical outlet ends, and inlet ports facing the outlet port of the common gas pipe and the diameters of which are identical to the inside diameter of the annular sleeves. made of heat-resistant material.
  • the contact surfaces of the first end of each rod and of the annular body must be thermally insulated.
  • each pair of nozzles is produced in the annular body in the form of a rectilinear channel with frustoconical ends, the annular body having radial orifices, the axis of which passes on one side through the center of the corresponding rectilinear channel, is perpendicular to the axis of the latter and is in the same plane, and on the other side, is perpendicular to the axis of the outlet pipe of the corresponding common gas pipe, and is in a second plane, and finally the axis of these orifices is on the crossing of the first two planes, each rod being fixed to the annular body by a from its ends, by means of a spindle which is coated with a thermostable composite material, and arranged so as to ensure rotation in the radial orifice, coated with a heat-insulating layer; the layer of composite material of each spindle having an ejection orifice to ensure
  • the orientation means are produced in the form of pulse reaction motors, located in the annular body, in rows regular, each pulse motor nozzle being oriented perpendicular to the longitudinal axis of the gas pipe of the annular body, each row being formed by pulse motors of the same type and of the same dimensions.
  • This embodiment is characterized by the simplicity of mounting the orientation means in the annular body and makes it possible to ensure independence with respect to the operation of the aerodynamic control surfaces and the orientation means, by ensuring pitch control and cap.
  • the axes of the frustoconical outlet ends of the nozzles of these engines can be directed tangentially with respect to the annular body.
  • one can control the roll of the flying machine.
  • the flying object is a missile, launched vertically from a ground launching area or from a ship, but it is understood that this flying object can be launched (horizontally) from a flying carrier, and / or that this flying machine is not necessarily a missile, but can also be a drone, for example.
  • the missile launch and orientation system 1 (FIG. 1) comprises aerodynamic control surfaces 2 with their means drive (not shown) which are usually arranged inside the missile, the annular body 3 and the launching means (not shown in Figure 1).
  • the annular body 3 comprises orientation means comprising a gas generator 4 and nozzles 5 which are connected to it and which open to the external surface of the annular body 3 of the missile 1.
  • Inside the body of the missile 1 is located the cruising engine with the nozzle 6, coaxial with the annular body 3.
  • the internal surface of the annular body 3 has a conical shape and is covered with a composite heat-insulating material, for example containing carbon. It forms a nozzle section 7, the profile of which is the continuation of the profile of the nozzle 6 of the cruise engine 6 of the missile (as shown in FIG. 4).
  • the design of the annular body 3 allows its ejection of the missile 1 in flight, since it is fixed to the body of the missile 1 using explosive bolts 8 and pyro-pushers 9 ( Figure 4).
  • the launching means include a launch container 10, a pressure generator 11 and a protective shutter 12 ( Figure 4).
  • the launch container 10 has front and rear covers. Its internal volume has a cylindrical shape and has dimensions making it possible to accommodate the missile 1 with the control surfaces 2 folded (the upper part of the container with the front cover is not shown in the drawing).
  • the pressure generator 11 is located at the bottom of the launch container 10, closed by the removable rear cover 13.
  • the annular body 3 has, in its rear part, a peripheral valve 15, the outside diameter of which is equal to the diameter inside of the container 10.
  • the protective shutter 12, intended to be mounted in a leaktight manner (like a plug) in the nozzle section 7 of the annular body 3, has a convex shape and a conical lateral surface, the profile of which is the same as that of the inner surface of the nozzle section 7 with which this shutter is in contact.
  • the convex part of the shutter 12 is on the side of the smaller diameter (that is to say it is oriented towards the cruise engine of the missile).
  • the shutter can be either metallic, or a composite heat-insulating material, for example epoxy resin with a graphite additive.
  • the launch container 10 comprises, in the fixing zone of the annular body 3, facing the valve 15, a gas ejection orifice 16 (FIG. 5).
  • the dimensions of the ejection orifice 16 are chosen taking into account the flow rate of the jet which passes through the ejection orifice 16.
  • the front cover of the container 10 must be fragmentable at a given pressure, produced inside the container . To do this, it is made of a fragile polymer, for example of polyurethane foam of strictly defined thickness, and this cover is fixed in a hermetic manner on the container 10.
  • Each mode has its own design of the annular body 3 and its own method of operating the orientation equipment.
  • the nozzles 5 of the orientation means are located in the same plane, perpendicular to the longitudinal axis of the gas pipe 7 of the annular body 3 (see Figure 1, Figure 4, Figure 6 and Figure 7) , whereas in the second embodiment, they are located on several planes (cf. FIG. 8).
  • the orientation of the missile 1 is ensured in pitch, heading and roll.
  • the first embodiment of the system in turn assumes two variants.
  • the first variant is illustrated in FIGS. 1, 2 and 3, and the second variant in FIGS. 4, 6 and 7.
  • the two variants of the first embodiment include an annular gas generator 4 (for example, with solid fuel) , located in the annular body 3, in which the supply gas lines 17 are located, connecting the gas generator 4 to the nozzles 5 (cf. FIG. 1 and FIG. 4).
  • the nozzles 5 are identical and grouped in pairs, the axes of which are situated in the same plane, each pair having its own gas supply 17 (cf. FIG. 2 and FIG. 6).
  • the nozzles 5 of each pair are oriented in opposition to each other and are connected at one end to the corresponding rod 18.
  • the number of rods 18 is identical to the number of control surfaces 2, which can be four in number.
  • Each rod 18 is fixed to the annular body 3 and its second end is connected to its control surface 2 by means of a "V" shaped fork 19 (cf. FIG. 1 and FIG. 4) fixed by hinges on the rod 18, encircling the rear edge of the control surface 2 and pushed towards the control surface by a spring (the latter is not shown in the drawing).
  • This spring ensures the interaction of the couple (fork 19 - control surface 2). As will be seen in what is explained below, this ensures the possibility of a joint rotation of the rods 18 with the control surfaces 2, which results in the required distribution of the gas jet which is constantly ejected from each pipe. gas 17, for each pair of nozzles 5.
  • the rods 18 are fixed in their middle part on the annular body by means of their axes of rotation 20 (cf. FIG. 1) each rod 18 enters contact with the annular body 3 through its first end, which comprises the pair of nozzles 5 produced in the form of bent channels ending in coaxial frustoconical ends, oriented in opposite directions (cf. FIG. 3).
  • the inlet ports of these bent channels open onto the outlet port of their common gas lines 17. In the area of these ports, the annular body and the end of the rod 18, which is in contact with it.
  • thermo-insulating plates 21 and 22 in composite material with an additive to graphite, the plates 21 and 22 are essential to prevent the erosion of the surfaces of contact under the influence of hot gas which passes by the orifices of the pair "rod 18 - annular body 3".
  • the pads 21 and 22 provide this protective function in combination with thermostable sleeves 23, which can be made from the same composite material.
  • Each sleeve 23 is inserted into the corresponding nozzle section 7, with the possibility of a longitudinal displacement, that is to say that the outside diameter of the sleeve 23 is practically equal to the diameter of the gas pipe 17.
  • the inside diameter of the sleeve 23 must be equal to the diameters of the orifices for receiving the nozzles with bent channels 5. Otherwise, as follows from what is explained below, the principle of operation of this sub-assembly cannot be satisfactorily ensured.
  • the second variant of the first embodiment of the system of the invention comprises rotary distributors which control the arrival of the gas in the pairs of nozzles 5, located, as can be seen in FIGS. 6 and 7, directly at the inside the annular body 3 in the form of rectilinear channels with frustoconical ends oriented in opposite directions.
  • the rotary distributors are produced in the following manner: in the annular body 3, radial holes 24 are drilled (FIG.
  • each radial orifice 24 is disposed a rotary pin 25 which is rigidly connected using, for example, a bolt 26 (see Figure 6) at the first end of the rod 18 (see Figure 4).
  • Each pin 25, as well as the contact surface of the radial orifice 24 in the annular body 3, is covered with a heat-insulating layer 27, 28 of composite material such as that mentioned above.
  • the functional role of the heat-insulating layers 27 and 28 is the same as that of the plates 21 and 22 in the first variant of the first embodiment, namely; prevent deterioration of the contact surfaces of the moving torque of the parts.
  • a groove 27A is practiced, the dimensions of which condition the distribution of the gas jet from the gas line 17 between the nozzles 5 of each pair .
  • the dimensions of the groove 27A are chosen so as to ensure a progressive modification during the rotation of the spindle 25 from an extreme position, for which the gas can arrive from the common channel 17 only towards one of the nozzles 5, towards a position for which the gas is equally distributed between the two nozzles 5 of the torque.
  • the depth of this groove 27A formed in the layer 27 is determined by the minimum thickness of this heat-insulating layer, necessary for the protection of the pin 25.
  • the second embodiment of the system of the invention provides for the use, as means of orientation, of standard components: impulsive jet engines operating with solid fuel, produced in a known manner in oneself.
  • a large quantity of these pulse motors (for example, several tens) are arranged on the periphery of the annular body 3, in regular rows 29-32, distributed over its height.
  • Each 29k-32k pulse motor is fixed in a housing made in the annular body 3, its nozzle being oriented perpendicular to the longitudinal axis of the nozzle section 7.
  • Each row 29-32 is formed by identical pulse motors, that is to say by motors of the same dimensions and of the same type in the row considered. From one row to another, the dimensions and types of the motors can be different or identical.
  • such use of standard pulse motors provides missile control only in pitch and heading (yaw).
  • the frustoconical end pieces of these nozzles are oriented in such a way that their axes are directed tangentially with respect to the annular body 3.
  • This orientation of the end pieces must be practiced, at a minimum, for the pulse motors of the row of motors of lower power, for example row 29.
  • the pulse motors of row of motors of lower power for example row 29.
  • half of the pulse motors of row 29 must have their nozzle oriented in the same direction (for example, clockwise shows around the axis of the nozzle section 7), while the second half must be oriented in the other direction (anti-clockwise).
  • rows 29 and 30 must be composed of pulse motors of the same type. It is preferable to use, to control the roll of the missile, the motors with lower power pulse. Indeed, to control the roll of missile 1, it is not necessary to create reactive forces as large as those which are necessary to control pitch and heading.
  • the missile launch and orientation system works as follows.
  • Missile 1 for example of the "ground-air" type with the annular body 3, produced either in accordance with FIG. 1 (see also FIGS. 2 and 3), or in accordance with FIG. 4 (see also FIGS. 6 and 7 ), or in accordance with FIG. 8, is disposed in the vertical launch container 10, the rear cover 13 of which is removed (cf. FIG. 4 and FIG. 8).
  • the missile 1 is then in a transport state (that is to say with the control surfaces 2 folded) while the protective shutter 12 is applied in a sealed manner on the nozzle section 7 of the annular body 3
  • the annular body 3 is connected to the support 14 using explosive bolts, after which a pressure generator 11 is placed in the container 10, and the rear cover 13 is closed at the front, the container 10 being hermetically closed. with the front cover.
  • the system of the invention is assembled and ready to operate.
  • Part of the gas is ejected through the orifice 16 (cf. FIG. 5) towards the hermetic upper cavity of the container 10.
  • the pressure under the front cover of the container 10 reaches a critical level, destruction of the front cover occurs. and ejecting debris outward.
  • the bolts that hold the missile on the support 14 and the valve 15 of the missile explode, sliding along the surface inner cylindrical guide of the container 10 closes the orifice 16, and the missile shoots upwards and is ejected at the required height (which can reach for example 40m), necessary for the execution of the maneuver for the orientation of the missile and the starting the cruise engine in difficult launch conditions.
  • the maneuvers are carried out for the orientation of the missile, i.e. the control pitch, heading and roll.
  • the execution of these maneuvers is carried out differently depending on the embodiment of the means for orienting the annular body 3.
  • the angular position of the control surfaces 2 controls the angular position of the corresponding rod 18, and the distribution of the gas jet between the nozzles 5 of the corresponding torque is carried out in proportion to the angular position of the rod 18, and thereby creates reaction forces of the same sign as in the aerodynamic planes of the control surface 2, ensuring the control of the missile in pitch, heading and roll.
  • the principle of creation of the steering reaction forces is similar to that which is mentioned above.
  • the difference lies only in the fact that in the second variant, the rotation of the rod 18 is controlled by the rotation of the control surface 2, which causes the rotation of the spindle 25 (cf. FIG. 7).
  • the angular position of the spindle 25 determines the quantity of gas which arrives in each nozzle 5 of the couple, and therefore the value of the result of the reaction forces in the couple of nozzles.
  • the principle of creation of the reaction forces which control the missile 1 is a little different from that which is described above.
  • the orientation of the missile 1 is carried out without the participation of the aerodynamic control surfaces 2, thanks to the start-up at a given instant of the pulse reaction engines, controlled for example directly by the computer of the electronic block of the missile.
  • the rocking of the missile in pitch and heading is ensured by the start-up of the most powerful impulse motors of rows 31-32, whose nozzles produce reaction forces oriented in a radial manner.
  • the direction of the rocking plane of the missile is determined by the low-power pulse motors of rows 29 and 30, whose nozzles produce reaction forces tangent to the annular body 3.
  • the cruise engine of the missile starts.
  • the gases produced during the operation of the cruise engine easily eject the protective shutter 12 (cf. FIGS. 1, 4 and 8) and after that, are freely ejected by the nozzle section 7 of the annular body 3, increasing the speed of the missile.
  • the profile of the nozzle section 7 is in continuity with the profile of the nozzle 6 of the cruise engine, the divergence of the nozzle of the cruise engine is optimized, which increases the momentum of the reaction force of the cruise engine in operation and compensates for a possible loss of speed, due to the presence of the inert mass of the annular body 3, representing the orientation means, which has already fulfilled its role.
  • the missile carries the inert mass far enough from the launch pad without additional energy consumption and, if necessary can eject it from the missile at a given time and in a given place.
  • the present invention allows, with a minimum of energy consumption, the interception of a target appeared suddenly near the launch pad, located in a difficult environment, and at the same time reduce to a minimum the harmful impact of the missile launch on the launching area by eliminating the need to eject the inertial mass from the orientation means after the performance of their function.
  • the invention can be applied to both large and small missiles.
  • the invention allows, with minimal modification of existing missiles with inclined launch, to give them all the qualities mentioned above.
  • the three modifications proposed in the particular cases of implementation of the launch and orientation control system of the missile are, from the point of view of the qualitative parameters, equivalent. The choice of one or the other is determined by the specificity of the missile which will have to use them. The means used in given circumstances may be less appropriate in other conditions.

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Abstract

A launching system, particularly a missile launching and steering system, whereby angle-launched missiles may be modernised by converting them into controlled missiles for circular defence, and ejection of the passive mass on the launch pad is prevented. For this purpose, the launching system includes a driven launching means (10) for the aerodynamic control surfaces (2) and a steering means comprising at least one gas generator (4) and nozzles (5) joined thereto. The steering means consists of an annular body (3) rigidly connected to the rear portion of the missile body. The inner surface of the annular body is tapered and covered with a thermal insulating material to form a gas ducting (7) with a profile in line with that of the sustainer motor nozzle (6) of the missile.

Description

La présente invention concerne des systèmes de lancement d'engins volants, et notamment des systèmes de lancement et d'orientation de missiles. Elle peut trouver son utilisation pour des missiles de petites ou grandes dimensions, du type "sol-air" ou "air-air" ou "sol-sol".The present invention relates to launching systems for flying machines, and in particular to missile launching and orientation systems. It can find its use for missiles of small or large dimensions, of the "ground-air" or "air-air" or "ground-ground" type.

Tout système de lancement et d'orientation d'engins volants comprend des moyens électroniques de commande et d'alimentation, ainsi que des moyens nécessaires à la mise en oeuvre du lancement et de l'orientation (moyens mécaniques, pyrotechniques,...) sous la commande desdits moyens électroniques.Any system for launching and orienting flying machines comprises electronic control and supply means, as well as means necessary for the implementation of launching and orientation (mechanical, pyrotechnic means, etc.) under the control of said electronic means.

On connaît un système de lancement et d'orientation de missile d'après les brevets USA n° 3 286 956 et US-A-2 995 319 (base pour le préambule de la revendication 1), qui comprend des moyens de lancement, des gouvernes aérodynamiques avec leurs dispositifs d'entraînement, ainsi que des moyens d'orientation comprenant essentiellement un générateur de gaz et des tuyères qui lui sont reliées.A missile launch and orientation system is known from US Pat. Nos. 3,286,956 and US-A-2,995,319 (base for the preamble of claim 1), which comprises launching means, aerodynamic control surfaces with their driving devices, as well as orientation means essentially comprising a gas generator and nozzles connected thereto.

Dans ce système, l'arrivée des gaz chauds s'effectue à partir du générateur de gaz qui est situé dans le corps du missile, à travers les axes de rotation des gouvernes, vers des tuyères situées dans la partie arrière des gouvernes et formant des jets réactifs dirigés parallèlement aux plans des gouvernes. Dans le monde, il existe un parc considérable de missiles, nécessitant une modernisation parce que ces missiles ne permettent pas d'assurer une défense omnidirectionnelle (c'est à dire d'intercepter une cible qui apparaît subitement de n'importe quelle direction par rapport à l'objectif à défendre). Théoriquement, il est possible de moderniser un missile à support de lancement incliné en le dotant du système connu, mentionné ci-dessus.In this system, the hot gases arrive from the gas generator which is located in the body of the missile, through the axes of rotation of the control surfaces, towards nozzles located in the rear part of the control surfaces and forming reactive jets directed parallel to the control surfaces. In the world, there is a considerable fleet of missiles, requiring modernization because these missiles do not provide an omnidirectional defense (i.e. to intercept a target which suddenly appears from any direction relative to to the objective to be defended). Theoretically, it is possible to modernize a missile with an inclined launch support by equipping it with the known system mentioned above.

Néanmoins, cela impliquerait de telles modifications dans la conception du missile que cela serait trop onéreux. En outre, le système de lancement et d'orientation considéré n'utilise pas entièrement l'énergie du jet réactif, jet qui est parallèle au plan des gouvernes, ce qui diminue la vitesse angulaire du missile lors de son changement de direction vers la cible.However, this would imply such modifications in the design of the missile that it would be too expensive. In addition, the launch and orientation system under consideration does not fully use the energy of the reactive jet, a jet which is parallel to the plane of the control surfaces, which decreases the angular speed of the missile when it changes direction towards the target. .

On connaît un système de lancement et d'orientation de missiles (Brevet international WO 94/10527) qui comprend des moyens de lancement, des gouvernes aérodynamiques avec leurs moyens d'entraînement, et des moyens d'orientation comportant des générateurs de gaz ainsi que des tuyères qui leur sont raccordées. Dans certains modes de réalisation, ce système connu comprend un générateur de gaz qui est relié par l'intermédiaire de conduites de gaz à des couples de tuyères ; chaque couple est formé de deux tuyères identiques, orientées dans des directions opposées dont les orifices d'admission donnent sur l'orifice de sortie de leur conduite de gaz commune, et dont les diamètres sont identiques à celui de l'orifice de sortie de la conduite de gaz.A system for launching and orienting missiles is known (International Patent WO 94/10527) which comprises launching means, aerodynamic control surfaces with their drive means, and orientation means comprising gas generators as well as nozzles connected to them. In certain embodiments, this known system comprises a gas generator which is connected via gas pipes to pairs of nozzles; each pair is formed by two identical nozzles, oriented in opposite directions, the inlet ports of which face the outlet of their common gas pipe, and the diameters of which are identical to that of the outlet of the gas line.

Ce système connu assure la possibilité d'un virage rapide du missile en direction de la cible, grâce au jet réactif éjecté de chaque couple de tuyères, et perpendiculaire au plan des gouvernes.This known system ensures the possibility of a rapid turn of the missile towards the target, thanks to the reactive jet ejected from each pair of nozzles, and perpendicular to the plane of the control surfaces.

Néanmoins, ainsi que dans le cas du système du susdit brevet US, dans ce brevet WO les moyens d'orientation forment un bloc en commun avec les moyens d'entraînement des gouvernes, ce qui est difficile à intégrer dans la conception des missiles de faibles dimensions sans que cela dégrade leurs propriété aérodynamiques. En outre, cela exclut la possibilité de largage, après le virage du missile dans la direction requise, de la masse inerte que représentent les moyens d'orientation. Ce système pourrait être utilisé également pour la modernisation mentionnée ci-dessus des missiles à lancement incliné.However, as in the case of the above-mentioned US patent system, in this WO patent the orientation means form a block in common with the drive means for the control surfaces, which is difficult to integrate into the design of weak missiles. dimensions without degrading their aerodynamic properties. In addition, this excludes the possibility of dropping, after the missile has turned in the required direction, the inert mass represented by the orientation means. This system could also be used for the above-mentioned modernization of inclined-launch missiles.

Le système de commande décrit dans l'article de Roger P.Berry, "Development of an orientation control system of the advanced kinetic energy missile" (ADKEM), AIAA-92-2763, comprend également des moyens de lancement, des gouvernes aérodynamiques avec entraînement, ainsi que des moyens d'orientation destinés à être implantés dans la partie arrière du missile, et dont la réalisation est basée sur des générateurs de gaz reliés à des tuyères.The control system described in the article by Roger P. Berry, "Development of an orientation control system of the advanced kinetic energy missile" (ADKEM), AIAA-92-2763, also includes launching means, aerodynamic control surfaces with drive, as well as orientation means intended to be installed in the rear part of the missile, and the production of which is based on gas generators connected to nozzles.

Le système décrit dans cet article peut être adapté à des missiles à lancement incliné (pour effectuer la modernisation mentionnée ci-dessus) sans modification considérable de ces missiles. Ce système prévoit le largage de la masse inerte des moyens d'orientation après exécution de leur fonction. Néanmoins, la complexité du système, l'encombrement important des moyens d'orientation qui sont prévus exclusivement pour l'utilisation de combustibles liquides hautement toxiques (hydrazine), rendent très difficile la mise en oeuvre de ce système.The system described in this article can be adapted to missiles with inclined launch (to carry out the modernization mentioned above) without considerable modification of these missiles. This system provides for the release of the inert mass of the orientation means after the performance of their function. However, the complexity of the system, the large size orientation means which are provided exclusively for the use of highly toxic liquid fuels (hydrazine), make the implementation of this system very difficult.

Du fait que lesdits moyens d'orientation sont situés sur le trajet des gaz éjectés par les tuyères des moteurs de croisière du missile, il est nécessaire de prévoir le largage des moyens d'orientation tout de suite après le virage en direction de la cible. Par ailleurs, ce largage, doit être effectué immédiatement avant l'allumage des moteurs de croisière, c'est à dire au-dessus de l'aire de lancement, ce qui complique l'exécution des actions militaires et par ailleurs est dangereux pour l'objectif à défendre.Because said orientation means are located on the path of the gases ejected by the nozzles of the cruise engines of the missile, it is necessary to provide for the release of the orientation means immediately after the turn towards the target. Furthermore, this release must be carried out immediately before the ignition of the cruise engines, that is to say above the launch pad, which complicates the execution of military actions and is also dangerous for the objective to defend.

Aucun des systèmes de lancement et d'orientation de missiles mentionnés ci-dessus ne permet d'assurer l'interception d'un cible proche dans les conditions difficiles d'un départ vertical, par exemple à partir de l'aire située dans un massif forestier. Ceci est tout d'abord lié à la réalisation des moyens de lancement de ces systèmes, moyens qui ne permettent pas d'atteindre rapidement une hauteur de l'ordre de 40 m nécessaire pour accomplir parfaitement les manoeuvres d'orientation vers la cible et l'allumage du moteur de croisière.None of the missile launch and orientation systems mentioned above can ensure the interception of a close target in the difficult conditions of a vertical departure, for example from the area located in a massif forest. This is first of all linked to the production of the means for launching these systems, means which do not allow a height of the order of 40 m to be quickly reached, necessary for perfectly performing the maneuvers of orientation towards the target and the ignition of the cruise engine.

Le problème principal que doit résoudre la présente invention est la réalisation d'un système universel de lancement et d'orientation du missile, qu'il serait possible d'associer aussi bien à des missiles de grandes que de faibles dimensions, permettant le largage de la masse inerte des moyens d'orientation suffisamment loin de l'aire de lancement. Ce système doit être le moins onéreux possible et doit pouvoir être utilisé pour tous les missiles à départ incliné, et doit pouvoir assurer une défense omnidirectionnelle.The main problem to be solved by the present invention is the realization of a universal missile launch and orientation system, which it would be possible to combine both large and small missiles, allowing the launching of the inert mass of the orientation means far enough from the launch pad. This system must be as inexpensive as possible and must be able to be used for all missiles with inclined departure, and must be able to provide omnidirectional defense.

Le système de lancement et d'orientation d'engins volants conforme à l'invention comprend des moyens de lancement, des gouvernes aérodynamiques avec leur entraînement et des moyens d'orientation, situés dans la partie arrière de l'engin volant et comportant au moins un générateur de gaz et des tuyères qui lui sont reliées, et ce système est caractérisé par le fait qu'il comporte un corps annulaire relié de façon rigide au corps de l'engin volant, les moyens d'orientation étant situés dans le corps annulaire, la surface interne du corps annulaire ayant une forme en tronc de cône et étant revêtue d'un matériau thermo-isolant, formant une section de tuyère dont le profil est dans la continuité du profil de la tuyère du moteur de croisière de l'engin volant.The launching and orientation system of flying machines according to the invention comprises launching means, aerodynamic control surfaces with their drive and orientation means, located in the rear part of the flying vehicle and comprising at least a gas generator and nozzles connected to it, and this system is characterized in that it comprises an annular body rigidly connected to the body of the flying object, the orientation means being located in the annular body , the internal surface of the annular body having a frusto-conical shape and being coated with a heat-insulating material, forming a nozzle section whose profile is in continuity with the profile of the nozzle of the cruising engine of the flying machine.

Le corps annulaire peut comporter des moyens assurant son éjection par l'engin volant au cours du vol, ce qui permet d'optimiser le bilan énergétique et de larguer entièrement la masse inerte que représentent les moyens d'orientation après leur utilisation, à un instant choisi, en dehors de la zone de l'aire de lancement.The annular body may include means ensuring its ejection by the flying object during the flight, which makes it possible to optimize the energy balance and to entirely release the inert mass which the orientation means represent after their use, at an instant. chosen, outside the launch pad area.

Selon un mode de réalisation, les tuyères des moyens d'orientation sont situées dans un même plan, perpendiculaire à l'axe longitudinal de la section de tuyère. Ceci assure une utilisation optimale de l'énergie des jets réactifs lors de l'orientation de l'engin volant, et, par conséquent, permet l'interception de la cible à proximité de l'aire de lancement.According to one embodiment, the nozzles of the orientation means are located in the same plane, perpendicular to the longitudinal axis of the nozzle section. This ensures optimal use of the energy of the reactive jets during the orientation of the flying object, and, consequently, allows the interception of the target near the launch pad.

Dans le cas d'un lancement vertical ou incliné, les moyens de lancement sont réalisés sous forme d'un conteneur de lancement avec des couvercles avant et arrière, dont le volume intérieur a une forme cylindrique et est destiné à recevoir l'engin volant, le générateur de pression étant situé au fond du conteneur, fermé par un couvercle arrière et par un obturateur de protection ayant une surface latérale tronconique, dont le profil reproduit au moins certaines parties de la surface de la section de tuyère du corps annulaire. La partie arrière du corps annulaire comporte un clapet périphérique, dont le diamètre extérieur est égal au diamètre intérieur du conteneur. Le conteneur comporte un support sur lequel sont fixés des éléments fragilisés destinés à la fixation du corps annulaire au-dessus des orifices de sortie du générateur de pression . Ceci assure le lancement de l'engin volant à partir du conteneur de lancement à l'aide du générateur de pression, ce qui permet d'intercepter une cible qui apparaît de façon soudaine à proximité de l'aire de lancement, dans des conditions de lancement difficiles (par exemple, au milieu d'un massif forestier ou sur le pont d'un navire comportant des superstructures élevées).In the case of a vertical or inclined launch, the launch means are produced in the form of a launch container with front and rear covers, the internal volume of which has a cylindrical shape and is intended to receive the flying machine, the pressure generator being located at the bottom of the container, closed by a rear cover and by a protective shutter having a frustoconical lateral surface, the profile of which reproduces at least certain parts of the surface of the nozzle section of the annular body. The rear part of the annular body has a peripheral valve, the outside diameter of which is equal to the inside diameter of the container. The container has a support on which are weakened elements intended for fixing the annular body above the outlet openings of the pressure generator. This ensures the launch of the flying object from the launch container using the pressure generator, which makes it possible to intercept a target that suddenly appears near the launch area, under conditions of difficult launching (for example, in the middle of a forest massif or on the deck of a ship with high superstructures).

Selon un mode de réalisation préféré de l'invention, l'obturateur de protection a une forme convexe orientée vers le moteur de croisière. Cette réalisation de l'obturateur permet d'assurer, ainsi que décrit ci-dessous, une fiabilité et une efficacité maximales de son fonctionnement dans le système de lancement.According to a preferred embodiment of the invention, the protective shutter has a convex shape oriented towards the cruising engine. This embodiment of the shutter ensures, as described below, maximum reliability and efficiency of its operation in the launch system.

Le conteneur de lancement peut comporter, dans la partie de fixation du corps annulaire, un orifice d'éjection dont les dimensions sont choisies compte tenu du débit de gaz passant par le jeu qui est formé autour du clapet du corps annulaire. Le couvercle avant du conteneur est réalisé de façon a être fragmenté pour une pression donnée se développant à l'intérieur du conteneur. Ces caractéristiques assurent une auto-éjection en temps voulu du couvercle avant du conteneur de lancement, avec une consommation d'énergie minimale, immédiatement avant le lancement de l'engin volant.The launch container may include, in the fixing part of the annular body, an ejection orifice, the dimensions of which are chosen taking into account the gas flow rate passing through the clearance which is formed around the valve of the annular body. The front cover of the container is made so as to be fragmented for a given pressure developing inside the container. These characteristics ensure timely self-ejection of the front cover of the launch container, with minimum energy consumption, immediately before the launch of the flying object.

Dans le premier mode de réalisation de l'invention, le système de lancement et d'orientation de l'engin volant peut être doté de tringles fixées sur le corps annulaire. Le générateur de gaz est également annulaire et relié aux tuyères des moyens d'orientation par des conduites de gaz formées dans le corps annulaire, les tuyères étant toutes identiques, groupées par deux dans le même plan. Les tuyères de chaque couple sont orientées de façons opposées et reliées mécaniquement à une extrémité de la tringle correspondante, ce qui assure la répartition du jet de gaz entre elles à partir de la conduite de gaz commune du corps annulaire. Chaque tringle est reliée par son autre extrémité à une gouverne correspondante assurant ainsi la possibilité d'une rotation conjointe. Par conséquent la rotation des gouvernes aérodynamiques et des moyens d'orientation est commandée par un moyen d'entraînement unique.In the first embodiment of the invention, the launching and orientation system of the flying machine can be provided with rods fixed to the annular body. The gas generator is also annular and connected to the nozzles of the orientation means by gas pipes formed in the annular body, the nozzles being all identical, grouped in pairs in the same plane. The nozzles of each pair are oriented in opposite ways and mechanically connected to one end of the corresponding rod, which ensures the distribution of the gas jet between them from the common gas pipe of the annular body. Each rod is connected by its other end to a corresponding control surface thus ensuring the possibility of a joint rotation. Consequently, the rotation of the aerodynamic control surfaces and of the orientation means is controlled by a single drive means.

La présente invention prévoit deux variantes du premier mode de réalisation du système de lancement et d'orientation d'engin volant. Selon la première variante, le système de commande est doté de manchons annulaires en matériau thermorésistant situés près de l'extrémité de sortie de chaque conduite de gaz correspondante, ces manchons pouvant se déplacer longitudinalement. Chaque tringle est fixée au corps annulaire dans sa partie médiane par son axe de rotation. Chaque couple de tuyères est réalisé sous forme de conduites coudées avec des extrémités de sortie tronconiques, et des orifices d'admission faisant face à l'orifice de sortie de la conduite de gaz commune et dont les diamètres sont identiques au diamètre intérieur des manchons annulaires en matériau thermorésistant. Les surfaces de contact de la première extrémité de chaque tringle et du corps annulaire doivent être thermo-isolées.The present invention provides two variants of the first embodiment of the launching and orientation system of the flying machine. According to the first variant, the control system is provided with annular sleeves made of heat-resistant material located near the outlet end of each corresponding gas pipe, these sleeves being able to move longitudinally. Each rod is fixed to the annular body in its middle part by its axis of rotation. Each pair of nozzles is produced in the form of bent pipes with frustoconical outlet ends, and inlet ports facing the outlet port of the common gas pipe and the diameters of which are identical to the inside diameter of the annular sleeves. made of heat-resistant material. The contact surfaces of the first end of each rod and of the annular body must be thermally insulated.

Dans la deuxième variante du premier mode de réalisation du système de lancement et d'orientation d'engins volants de l'invention, chaque couple de tuyères est réalisé dans le corps annulaire sous forme d'un canal rectiligne à embouts tronconiques, le corps annulaire comportant des orifices radiaux, dont l'axe passe d'un côté par le centre du canal rectiligne correspondant, est perpendiculaire à l'axe de ce dernier et se trouve dans un même plan, et de l'autre côté, est perpendiculaire à l'axe du tuyau de sortie de la conduite de gaz commune correspondante, et se trouve dans un deuxième plan, et enfin l'axe de ces orifices se trouve sur le croisement des deux premiers plans, chaque tringle étant fixée sur le corps annulaire par une de ses extrémités, par l'intermédiaire d'une broche qui est revêtue d'un matériau thermostable composite, et disposée de façon à assurer la rotation dans l'orifice radial, revêtu d'une couche thermo-isolante ; la couche de matériau composite de chaque broche comportant un orifice d'éjection pour assurer la répartition du jet de gaz entre les tuyère du couple.In the second variant of the first embodiment of the launching and orientation system of flying machines of the invention, each pair of nozzles is produced in the annular body in the form of a rectilinear channel with frustoconical ends, the annular body having radial orifices, the axis of which passes on one side through the center of the corresponding rectilinear channel, is perpendicular to the axis of the latter and is in the same plane, and on the other side, is perpendicular to the axis of the outlet pipe of the corresponding common gas pipe, and is in a second plane, and finally the axis of these orifices is on the crossing of the first two planes, each rod being fixed to the annular body by a from its ends, by means of a spindle which is coated with a thermostable composite material, and arranged so as to ensure rotation in the radial orifice, coated with a heat-insulating layer; the layer of composite material of each spindle having an ejection orifice to ensure the distribution of the gas jet between the nozzles of the torque.

Ces deux variantes du premier mode de réalisation du système de lancement et d'orientation de l'engin volant sont compactes, ont une technologie équivalente, et sont caractérisées par une grande fiabilité du fonctionnement de matériel d'orientation par l'entraînement des gouvernes aérodynamiques.These two variants of the first embodiment of the launching and orientation system of the flying machine are compact, have equivalent technology, and are characterized by a high reliability of the functioning of orientation equipment by driving the aerodynamic control surfaces. .

Dans le deuxième mode de réalisation du système de commande de lancement et d'orientation de l'engin volant selon l'invention, les moyens d'orientation sont réalisés sous forme de moteurs à réaction par impulsions, situés dans le corps annulaire, en rangées régulières, chaque tuyère de moteur à impulsion étant orientée perpendiculairement à l'axe longitudinal de la conduite de gaz du corps annulaire, chaque rangée étant formée par des moteurs à impulsion d'un même type et de mêmes dimensions.In the second embodiment of the launching and orientation control system of the flying object according to the invention, the orientation means are produced in the form of pulse reaction motors, located in the annular body, in rows regular, each pulse motor nozzle being oriented perpendicular to the longitudinal axis of the gas pipe of the annular body, each row being formed by pulse motors of the same type and of the same dimensions.

Ce mode de réalisation est caractérisé par la simplicité de montage des moyens d'orientation dans le corps annulaire et permet d'assurer l'indépendance par rapport au fonctionnement des gouvernes aérodynamiques et des moyens d'orientation, en assurant le contrôle de tangage et de cap.This embodiment is characterized by the simplicity of mounting the orientation means in the annular body and makes it possible to ensure independence with respect to the operation of the aerodynamic control surfaces and the orientation means, by ensuring pitch control and cap.

Dans le deuxième mode de réalisation du système de lancement et d'orientation de l'engin volant, au moins les moteurs à impulsion de plus faible puissance forment une rangée, les axes des extrémités tronconiques de sortie des tuyères de ces moteurs peuvent être dirigés tangentiellement par rapport au corps annulaire. Ainsi, on peut contrôler le roulis de l'engin volant.In the second embodiment of the launching and orientation system of the flying object, at least the more impulse motors low power form a row, the axes of the frustoconical outlet ends of the nozzles of these engines can be directed tangentially with respect to the annular body. Thus, one can control the roll of the flying machine.

La présente invention sera mieux comprise à la lecture de la description détaillée de plusieurs modes de réalisation, pris à titre d'exemples non limitatifs et illustrés par le dessin annexé, sur lequel :

  • la figure 1 est une vue latérale avec une coupe partielle du système de lancement et d'orientation du missile, illustrant la première variante du premier mode de réalisation de l'invention ;
  • la figure 2, est une coupe transversale du système de commande au niveau des tuyères du dispositif d'orientation, vue dans la coupe II-II, figure 1;
  • la figure 3, est une vue agrandie de la coupe partielle III de la figure 2 ;
  • la figure 4, est une vue latérale avec une coupe partielle du système de commande, illustrant la deuxième variante du premier mode de réalisation de l'invention ;
  • la figure 5, est une vue agrandie de la partie V de la figure 4 ;
  • la figure 6, est une vue en coupe transversale du corps annulaire du système de commande au niveau de l'axe horizontal des tuyères du matériel d'orientation, selon VI-VI de la figure 4 ;
  • la figure 7, est une vue agrandie de la section longitudinale du système de commande dans la partie des tuyères selon VII-VII de la figure 6; et
  • la figure 8, est une vue latérale avec coupe partielle du système de commande, illustrant le deuxième mode de réalisation de l'invention.
The present invention will be better understood on reading the detailed description of several embodiments, taken by way of nonlimiting examples and illustrated by the appended drawing, in which:
  • Figure 1 is a side view with a partial section of the missile launch and orientation system, illustrating the first variant of the first embodiment of the invention;
  • Figure 2 is a cross section of the control system at the nozzles of the orientation device, seen in section II-II, Figure 1;
  • Figure 3 is an enlarged view of the partial section III of Figure 2;
  • Figure 4 is a side view with a partial section of the control system, illustrating the second variant of the first embodiment of the invention;
  • Figure 5 is an enlarged view of part V of Figure 4;
  • Figure 6 is a cross-sectional view of the annular body of the control system at the horizontal axis of the nozzles of the orientation material, according to VI-VI of Figure 4;
  • Figure 7 is an enlarged view of the longitudinal section of the control system in the part of the nozzles according to VII-VII of Figure 6; and
  • Figure 8 is a side view with partial section of the control system, illustrating the second embodiment of the invention.

L'invention est décrite ci-dessous dans le cas où l'engin volant est un missile, lancé verticalement depuis une aire de lancement au sol ou depuis un navire, mais il est bien entendu que cet engin volant peut être lancé (horizontalement) depuis un porteur volant, et/ou que cet engin volant n'est pas nécessairement un missile, mais peut aussi être un drone, par exemple.The invention is described below in the case where the flying object is a missile, launched vertically from a ground launching area or from a ship, but it is understood that this flying object can be launched (horizontally) from a flying carrier, and / or that this flying machine is not necessarily a missile, but can also be a drone, for example.

Le système de lancement et d'orientation du missile 1 (figure 1) comprend des gouvernes aérodynamiques 2 avec leurs moyens d'entraînement (non représentés) qui sont habituellement disposés à l'intérieur du missile, le corps annulaire 3 et les moyens de lancement (non représentés en figure 1). Le corps annulaire 3 comprend des moyens d'orientation comportant un générateur de gaz 4 et des tuyères 5 qui lui sont reliées et qui débouchent à la surface externe du corps annulaire 3 du missile 1. A l'intérieur du corps du missile 1 se trouve le moteur de croisière avec la tuyère 6, coaxial avec le corps annulaire 3. La surface interne du corps annulaire 3 a une forme conique et est recouverte d'un matériau thermo-isolant composite, contenant par exemple du carbone. Elle forme une section de tuyère 7, dont le profil est la continuation du profil de la tuyère 6 du moteur de croisière 6 du missile (comme représenté sur la figure 4).The missile launch and orientation system 1 (FIG. 1) comprises aerodynamic control surfaces 2 with their means drive (not shown) which are usually arranged inside the missile, the annular body 3 and the launching means (not shown in Figure 1). The annular body 3 comprises orientation means comprising a gas generator 4 and nozzles 5 which are connected to it and which open to the external surface of the annular body 3 of the missile 1. Inside the body of the missile 1 is located the cruising engine with the nozzle 6, coaxial with the annular body 3. The internal surface of the annular body 3 has a conical shape and is covered with a composite heat-insulating material, for example containing carbon. It forms a nozzle section 7, the profile of which is the continuation of the profile of the nozzle 6 of the cruise engine 6 of the missile (as shown in FIG. 4).

La conception du corps annulaire 3 permet son éjection du missile 1 en vol, étant donné qu'il est fixé sur le corps du missile 1 à l'aide de boulons explosifs 8 et de pyro-poussoirs 9 (figure 4).The design of the annular body 3 allows its ejection of the missile 1 in flight, since it is fixed to the body of the missile 1 using explosive bolts 8 and pyro-pushers 9 (Figure 4).

Les moyens de lancement comprennent un conteneur de lancement 10, un générateur de pression 11 et un obturateur de protection 12 (figure 4). Le conteneur de lancement 10 est doté de couvercles avant et arrière. Son volume intérieur a une forme cylindrique et a des dimensions permettant d'y loger le missile 1 avec les gouvernes 2 repliées (la partie supérieure du conteneur avec le couvercle avant n'est pas représentée sur le dessin). Le générateur de pression 11 est situé au fond du conteneur de lancement 10, fermé par le couvercle arrière amovible 13. Au fond du conteneur 10 se trouve le support 14, destiné à la fixation du corps annulaire 3, monté avec le missile 1 au-dessus du générateur 11. La fixation du corps annulaire 3 sur le support 14 est assurée par des éléments explosifs, par exemple des boulons explosifs. Afin d'assurer le glissement du corps annulaire 3 le long de la surface intérieure cylindrique de guidage de la cavité du conteneur 10, le corps annulaire 3 a, dans sa partie arrière, un clapet périphérique 15, dont le diamètre extérieur est égal au diamètre intérieur du conteneur 10. L'obturateur de protection 12, destiné à être monté d'une façon étanche (comme un bouchon) dans la section de tuyère 7 du corps annulaire 3, a une forme convexe et une surface latérale conique, dont le profil est le même que celui de la surface intérieure de la section de tuyère 7 avec laquelle cet obturateur est en contact. La partie convexe de l'obturateur 12 se trouve du côté du diamètre inférieur (c'est à dire qu'elle est orientée vers le moteur de croisière du missile). L'obturateur peut être soit métallique, soit en matériau thermo-isolant composite, par exemple en résine époxyde avec un additif au graphite.The launching means include a launch container 10, a pressure generator 11 and a protective shutter 12 (Figure 4). The launch container 10 has front and rear covers. Its internal volume has a cylindrical shape and has dimensions making it possible to accommodate the missile 1 with the control surfaces 2 folded (the upper part of the container with the front cover is not shown in the drawing). The pressure generator 11 is located at the bottom of the launch container 10, closed by the removable rear cover 13. At the bottom of the container 10 is the support 14, intended for fixing the annular body 3, mounted with the missile 1 au- above the generator 11. The fixing of the annular body 3 to the support 14 is ensured by explosive elements, for example explosive bolts. In order to ensure the sliding of the annular body 3 along the inner cylindrical surface for guiding the cavity of the container 10, the annular body 3 has, in its rear part, a peripheral valve 15, the outside diameter of which is equal to the diameter inside of the container 10. The protective shutter 12, intended to be mounted in a leaktight manner (like a plug) in the nozzle section 7 of the annular body 3, has a convex shape and a conical lateral surface, the profile of which is the same as that of the inner surface of the nozzle section 7 with which this shutter is in contact. The convex part of the shutter 12 is on the side of the smaller diameter (that is to say it is oriented towards the cruise engine of the missile). The shutter can be either metallic, or a composite heat-insulating material, for example epoxy resin with a graphite additive.

Le conteneur de lancement 10 comporte dans la zone de fixation du corps annulaire 3, face au clapet 15, un orifice d'éjection de gaz 16 (figure 5). Les dimensions de l'orifice d'éjection 16 sont choisies compte tenu du débit du jet qui passe par l'orifice d'éjection 16. Le couvercle avant du conteneur 10 doit être fragmentable à une pression donnée, produite à l'intérieur du conteneur. Pour ce faire, il est fabriqué en polymère fragile, par exemple en mousse de polyuréthane d'épaisseur strictement définie, et ce couvercle est fixé d'une façon hermétique sur le conteneur 10.The launch container 10 comprises, in the fixing zone of the annular body 3, facing the valve 15, a gas ejection orifice 16 (FIG. 5). The dimensions of the ejection orifice 16 are chosen taking into account the flow rate of the jet which passes through the ejection orifice 16. The front cover of the container 10 must be fragmentable at a given pressure, produced inside the container . To do this, it is made of a fragile polymer, for example of polyurethane foam of strictly defined thickness, and this cover is fixed in a hermetic manner on the container 10.

On décrit ici deux modes de réalisation de ce système de lancement et d'orientation de missile. Chaque mode a sa propre conception du corps annulaire 3 et son propre procédé de fonctionnement du matériel d'orientation. Dans le premier cas, les tuyères 5 des moyens d'orientation sont situées dans le même plan, perpendiculairement à l'axe longitudinal de la conduite de gaz 7 du corps annulaire 3 (cf figure 1, figure 4, figure 6 et figure 7), alors que dans le deuxième mode de réalisation, elles sont situées sur plusieurs plans (cf. figure 8). Néanmoins, dans les deux cas, ainsi qu'il s'ensuit de ce qui est exposé ci-dessous, l'orientation du missile 1 est assurée en tangage, en cap et en roulis.Two embodiments of this missile launch and orientation system are described here. Each mode has its own design of the annular body 3 and its own method of operating the orientation equipment. In the first case, the nozzles 5 of the orientation means are located in the same plane, perpendicular to the longitudinal axis of the gas pipe 7 of the annular body 3 (see Figure 1, Figure 4, Figure 6 and Figure 7) , whereas in the second embodiment, they are located on several planes (cf. FIG. 8). However, in both cases, as follows from what is explained below, the orientation of the missile 1 is ensured in pitch, heading and roll.

Le premier mode de réalisation du système suppose à son tour deux variantes. La première variante est illustrée par les figures 1, 2 et 3, et la deuxième variante par les figures 4, 6 et 7. Les deux variantes du premier mode de réalisation comportent un générateur de gaz 4 annulaire (par exemple, à combustible solide), se trouvant dans le corps annulaire 3, dans lequel se trouvent les conduites de gaz d'alimentation 17, raccordant le générateur de gaz 4 aux tuyères 5 (cf. figure 1 et figure 4). Les tuyères 5 sont identiques et groupées par couples, dont les axes sont situés dans un même plan, chaque couple ayant sa propre arrivée de gaz 17 (cf. figure 2 et figure 6).The first embodiment of the system in turn assumes two variants. The first variant is illustrated in FIGS. 1, 2 and 3, and the second variant in FIGS. 4, 6 and 7. The two variants of the first embodiment include an annular gas generator 4 (for example, with solid fuel) , located in the annular body 3, in which the supply gas lines 17 are located, connecting the gas generator 4 to the nozzles 5 (cf. FIG. 1 and FIG. 4). The nozzles 5 are identical and grouped in pairs, the axes of which are situated in the same plane, each pair having its own gas supply 17 (cf. FIG. 2 and FIG. 6).

Les tuyères 5 de chaque couple sont orientées en opposition l'une par rapport à l'autre et sont raccordées par une extrémité à la tringle correspondante 18. Le nombre de tringles 18 est identique au nombre de gouvernes 2, qui peuvent être au nombre de quatre. Chaque tringle 18 est fixée sur le corps annulaire 3 et sa deuxième extrémité est reliée à sa gouverne 2 par l'intermédiaire d'une fourchette en forme de "V" 19 (cf. figure 1 et figure 4) fixée par des charnières sur la tringle 18, ceinturant le rebord arrière de la gouverne 2 et poussée vers la gouverne par un ressort (ce dernier n'est pas représenté sur le dessin). Ce ressort assure l'interaction du couple (fourchette 19 - gouverne 2). Ainsi qu'on le verra dans ce qui est exposé ci-dessous, cela assure la possibilité d'une rotation conjointe des tringles 18 avec les gouvernes 2, ce qui entraîne la répartition requise du jet de gaz qui est éjecté en permanence de chaque conduite de gaz 17, pour chaque couple de tuyères 5.The nozzles 5 of each pair are oriented in opposition to each other and are connected at one end to the corresponding rod 18. The number of rods 18 is identical to the number of control surfaces 2, which can be four in number. Each rod 18 is fixed to the annular body 3 and its second end is connected to its control surface 2 by means of a "V" shaped fork 19 (cf. FIG. 1 and FIG. 4) fixed by hinges on the rod 18, encircling the rear edge of the control surface 2 and pushed towards the control surface by a spring (the latter is not shown in the drawing). This spring ensures the interaction of the couple (fork 19 - control surface 2). As will be seen in what is explained below, this ensures the possibility of a joint rotation of the rods 18 with the control surfaces 2, which results in the required distribution of the gas jet which is constantly ejected from each pipe. gas 17, for each pair of nozzles 5.

Pour la première variante du premier mode de réalisation du système de l'invention, les tringles 18 sont fixées dans leur partie médiane sur le corps annulaire par l'intermédiaire de leurs axes de rotation 20 (cf. figure 1) chaque tringle 18 entre en contact avec le corps annulaire 3 par sa première extrémité, qui comporte le couple des tuyères 5 réalisées sous forme de canaux coudés se terminant par des embouts tronconiques coaxiaux, orientés dans des directions opposées (cf. figure 3). Les orifices d'admission de ces canaux coudés débouchent sur l'orifice de sortie de leurs conduites de gaz communes 17. Dans la zone de ces orifices, le corps annulaire et l'extrémité de la tringle 18, qui est en contact avec celui-ci, sont protégés par des plaquettes thermo-isolantes 21 et 22, en matériau composite avec un additif au graphite, les plaquettes 21 et 22 sont indispensables pour prévenir l'érosion des surfaces de contact sous l'influence du gaz chaud qui passe par les orifices du couple "tringle 18 - corps annulaire 3". Les plaquettes 21 et 22 assurent cette fonction de protection en combinaison avec des manchons thermostables 23, qui peuvent être fabriqués à partir du même matériau composite. Chaque manchon 23 est inséré dans la section de tuyère correspondante 7, avec possibilité d'un déplacement longitudinal, c'est à dire que le diamètre extérieur du manchon 23 est pratiquement égal au diamètre de la conduite de gaz 17. Le diamètre intérieur du manchon 23 doit être égal aux diamètres des orifices de réception des tuyères à canaux coudés 5. Dans le cas contraire, comme il s'ensuit de ce qui est exposé ci-dessous, le principe de fonctionnement de ce sous-ensemble ne peut être assuré de façon satisfaisante.For the first variant of the first embodiment of the system of the invention, the rods 18 are fixed in their middle part on the annular body by means of their axes of rotation 20 (cf. FIG. 1) each rod 18 enters contact with the annular body 3 through its first end, which comprises the pair of nozzles 5 produced in the form of bent channels ending in coaxial frustoconical ends, oriented in opposite directions (cf. FIG. 3). The inlet ports of these bent channels open onto the outlet port of their common gas lines 17. In the area of these ports, the annular body and the end of the rod 18, which is in contact with it. Ci, are protected by thermo-insulating plates 21 and 22, in composite material with an additive to graphite, the plates 21 and 22 are essential to prevent the erosion of the surfaces of contact under the influence of hot gas which passes by the orifices of the pair "rod 18 - annular body 3". The pads 21 and 22 provide this protective function in combination with thermostable sleeves 23, which can be made from the same composite material. Each sleeve 23 is inserted into the corresponding nozzle section 7, with the possibility of a longitudinal displacement, that is to say that the outside diameter of the sleeve 23 is practically equal to the diameter of the gas pipe 17. The inside diameter of the sleeve 23 must be equal to the diameters of the orifices for receiving the nozzles with bent channels 5. Otherwise, as follows from what is explained below, the principle of operation of this sub-assembly cannot be satisfactorily ensured.

La deuxième variante du premier mode de réalisation du système de l'invention comporte des répartiteurs rotatifs qui commandent l'arrivée du gaz dans les couples de tuyères 5, situés, ainsi qu'on le voit sur les figures 6 et 7, directement à l'intérieur du corps annulaire 3 sous forme de canaux rectilignes avec des embouts tronconiques orientés dans des directions opposées. Les répartiteurs rotatifs sont réalisés de la façon suivante : dans le corps annulaire 3, on perce des orifices radiaux 24 (figure 7), dont les axes passent, d'une part, par le centre du canal rectiligne correspondant des tuyères 5 et qui est perpendiculaire à l'axe de ce canal rectiligne et se trouve dans le même plan, et d'autre part, ils sont perpendiculaires à l'axe de la conduite de gaz correspondante 17 et se trouvent dans un deuxième plan. En outre, ces axes se trouvent sur le croisement du premier et du deuxième plans. Dans chaque orifice radial 24 est disposée une broche rotative 25 qui est reliée rigidement à l'aide, par exemple, d'un boulon 26 (cf. figure 6) à la première extrémité de la tringle 18 (cf. figure 4). Chaque broche 25, ainsi que la surface de contact de l'orifice radial 24 dans le corps annulaire 3, est recouverte d'une couche thermo-isolante 27 , 28 en matériau composite tel que celui mentionné ci-dessus. Le rôle fonctionnel des couches thermo-isolantes 27 et 28 est le même que celui des plaquettes 21 et 22 dans la première variante du premier mode de réalisation, à savoir ; empêcher la détérioration des surfaces de contact du couple mobile des pièces. Sur une partie de la périphérie de la couche 27 de matériau composite, appliqué sur la broche 25, on pratique une saignée 27A dont les dimensions conditionnent la répartition du jet de gaz à partir de la conduite de gaz 17 entre les tuyères 5 de chaque couple. Les dimensions de la saignée 27A sont choisies de façon à assurer une modification progressive lors de la rotation de la broche 25 d'une position extrême, pour laquelle le gaz peut arriver du canal commun 17 uniquement vers l'une des tuyères 5, vers une position pour laquelle le gaz est équiréparti entre les deux tuyères 5 du couple. Bien entendu, il est nécessaire d'exclure la possibilité d'une coupure simultanée du débit de gaz vers les deux tuyères 5 du couple. La profondeur de cette saignée 27A pratiquée dans la couche 27 est déterminée par l'épaisseur minimum de cette couche thermo-isolante, nécessaire à la protection de la broche 25.The second variant of the first embodiment of the system of the invention comprises rotary distributors which control the arrival of the gas in the pairs of nozzles 5, located, as can be seen in FIGS. 6 and 7, directly at the inside the annular body 3 in the form of rectilinear channels with frustoconical ends oriented in opposite directions. The rotary distributors are produced in the following manner: in the annular body 3, radial holes 24 are drilled (FIG. 7), the axes of which pass, on the one hand, through the center of the corresponding rectilinear channel of the nozzles 5 and which is perpendicular to the axis of this rectilinear channel and is in the same plane, and on the other hand, they are perpendicular to the axis of the corresponding gas pipe 17 and are in a second plane. In addition, these axes are located on the intersection of the first and second planes. In each radial orifice 24 is disposed a rotary pin 25 which is rigidly connected using, for example, a bolt 26 (see Figure 6) at the first end of the rod 18 (see Figure 4). Each pin 25, as well as the contact surface of the radial orifice 24 in the annular body 3, is covered with a heat-insulating layer 27, 28 of composite material such as that mentioned above. The functional role of the heat-insulating layers 27 and 28 is the same as that of the plates 21 and 22 in the first variant of the first embodiment, namely; prevent deterioration of the contact surfaces of the moving torque of the parts. On a part of the periphery of the layer 27 of composite material, applied to the spindle 25, a groove 27A is practiced, the dimensions of which condition the distribution of the gas jet from the gas line 17 between the nozzles 5 of each pair . The dimensions of the groove 27A are chosen so as to ensure a progressive modification during the rotation of the spindle 25 from an extreme position, for which the gas can arrive from the common channel 17 only towards one of the nozzles 5, towards a position for which the gas is equally distributed between the two nozzles 5 of the torque. Of course, it is necessary to exclude the possibility of a simultaneous cut off of the gas flow to the two nozzles 5 of the torque. The depth of this groove 27A formed in the layer 27 is determined by the minimum thickness of this heat-insulating layer, necessary for the protection of the pin 25.

Le deuxième mode de réalisation du système de l'invention, illustré en figure 8, prévoit l'utilisation, en tant que moyens d'orientation, de composants standard : des moteurs à réaction impulsifs fonctionnant avec du combustible solide, réalisés de façon connue en soi. Une grande quantité de ces moteurs à impulsion (par exemple, plusieurs dizaines) sont disposés à la périphérie du corps annulaire 3, par rangées régulières 29-32, réparties sur sa hauteur. Chaque moteur à impulsion 29k-32k est fixé dans un logement pratiqué dans le corps annulaire 3, sa tuyère étant orientée perpendiculairement à l'axe longitudinal de la section de tuyère 7. Chaque rangée 29-32 est formée par des moteurs à impulsion identiques, c'est-à-dire par des moteurs de mêmes dimensions et de même type dans la rangée considérée. D'une rangée à l'autre, les dimensions et types des moteurs peuvent être différents ou bien identiques. Comme décrit ci-dessous, une telle utilisation de moteurs à impulsion standard assure la commande du missile uniquement en tangage et en cap (lacet).The second embodiment of the system of the invention, illustrated in FIG. 8, provides for the use, as means of orientation, of standard components: impulsive jet engines operating with solid fuel, produced in a known manner in oneself. A large quantity of these pulse motors (for example, several tens) are arranged on the periphery of the annular body 3, in regular rows 29-32, distributed over its height. Each 29k-32k pulse motor is fixed in a housing made in the annular body 3, its nozzle being oriented perpendicular to the longitudinal axis of the nozzle section 7. Each row 29-32 is formed by identical pulse motors, that is to say by motors of the same dimensions and of the same type in the row considered. From one row to another, the dimensions and types of the motors can be different or identical. As described below, such use of standard pulse motors provides missile control only in pitch and heading (yaw).

Afin d'assurer la commande du missile 1 en roulis, il est nécessaire de procéder à une petite modification des tuyères des moteurs à impulsion standard. A cet effet, on oriente les embouts tronconiques de sortie de ces tuyères de telle façon que leurs axes soient dirigés tangentiellement par rapport au corps annulaire 3. Cette orientation des embouts doit être pratiquée, au minimum, pour les moteurs à impulsion de la rangée de moteurs de plus faible puissance, par exemple la rangée 29. Il est évident que dans ce cas la moitié des moteurs à impulsion de la rangée 29 doivent avoir leur embout orienté dans le même sens (par exemple, dans le sens des aiguilles d'une montre autour de l'axe de la section de tuyère 7), alors que la deuxième moitié doit être orientée dans l'autre sens (dans le sens contraire des aiguilles d'une montre). Mais il est possible d'obtenir le même résultat en orientant tous les embouts d'une rangée dans le sens des aiguilles d'une montre (par exemple de la rangée 29) et en orientant dans le sens contraire des aiguilles d'une montre tous les moteurs à impulsion d'une autre rangée, (par exemple la rangée 30). Dans ce dernier cas, les rangées 29 et 30 doivent être composées de moteurs à impulsion du même type. Il est préférable d'utiliser, pour contrôler le roulis du missile, les moteurs à impulsion de plus faible puissance. En effet, pour contrôler le roulis du missile 1, il n'est pas nécessaire de créer des forces réactives aussi importantes que celles qui sont nécessaires pour contrôler le tangage et le cap.In order to ensure control of missile 1 in roll, it is necessary to make a small modification of the nozzles of standard pulse motors. For this purpose, the frustoconical end pieces of these nozzles are oriented in such a way that their axes are directed tangentially with respect to the annular body 3. This orientation of the end pieces must be practiced, at a minimum, for the pulse motors of the row of motors of lower power, for example row 29. It is obvious that in this case half of the pulse motors of row 29 must have their nozzle oriented in the same direction (for example, clockwise shows around the axis of the nozzle section 7), while the second half must be oriented in the other direction (anti-clockwise). But it is possible to obtain the same result by orienting all the end pieces of a row in a clockwise direction (for example of row 29) and by orienting all anticlockwise pulse motors from another row (for example row 30). In the latter case, rows 29 and 30 must be composed of pulse motors of the same type. It is preferable to use, to control the roll of the missile, the motors with lower power pulse. Indeed, to control the roll of missile 1, it is not necessary to create reactive forces as large as those which are necessary to control pitch and heading.

Le système de lancement et d'orientation de missile fonctionne de la façon suivante.The missile launch and orientation system works as follows.

Le missile 1, par exemple du type "sol-air" avec le corps annulaire 3, réalisé soit conformément à la figure 1 (voir aussi les figures 2 et 3), soit conformément à la figure 4 (voir aussi les figures 6 et 7), soit conformément à la figure 8, est disposé dans le conteneur de lancement vertical 10, dont le couvercle arrière 13 est démonté (cf. figure 4 et figure 8). Le missile 1 se trouve alors dans un état de transport (c'est-à-dire avec les gouvernes 2 repliées) alors que l'obturateur de protection 12 est appliqué d'une façon étanche sur la section de tuyère 7 du corps annulaire 3. Le corps annulaire 3 est relié au support 14 à l'aide de boulons explosifs, après quoi on dispose dans le conteneur 10 un générateur de pression 11, et on referme le couvercle arrière 13 à l'avant, le conteneur 10 étant fermé hermétiquement avec le couvercle avant. Le système de l'invention est monté et prêt à fonctionner.Missile 1, for example of the "ground-air" type with the annular body 3, produced either in accordance with FIG. 1 (see also FIGS. 2 and 3), or in accordance with FIG. 4 (see also FIGS. 6 and 7 ), or in accordance with FIG. 8, is disposed in the vertical launch container 10, the rear cover 13 of which is removed (cf. FIG. 4 and FIG. 8). The missile 1 is then in a transport state (that is to say with the control surfaces 2 folded) while the protective shutter 12 is applied in a sealed manner on the nozzle section 7 of the annular body 3 The annular body 3 is connected to the support 14 using explosive bolts, after which a pressure generator 11 is placed in the container 10, and the rear cover 13 is closed at the front, the container 10 being hermetically closed. with the front cover. The system of the invention is assembled and ready to operate.

Les gaz formés lors de l'inflammation de la charge du générateur de pression 11, créent au fond du conteneur 10 une surpression qui agit sur l'extrémité de la partie arrière du corps 3. L'obturateur 12, de ce fait s'enfonce davantage dans la section de tuyère 7 en protégeant le moteur de croisière du missile des gaz chauds du générateur 11, ce qui évite le risque d'une mise en route spontanée du moteur de croisière. Une partie des gaz est éjectée par l'orifice 16 (cf. figure 5) vers la cavité supérieure hermétique du conteneur 10. Dès que la pression sous le couvercle avant du conteneur 10 atteint un niveau critique, il se produit une destruction du couvercle avant et l'éjection des débris vers l'extérieur. Une fois que la pression dans l'espace clos du fond du conteneur atteint la valeur requise, il se produit l'explosion des boulons qui retiennent le missile sur le support 14, et le clapet 15 du missile, en glissant le long de la surface intérieure de guidage cylindrique du conteneur 10 obture l'orifice 16, et le missile s'élance vers le haut et est éjecté à la hauteur requise (qui peut atteindre par exemple 40m), nécessaire à l'exécution de la manoeuvre pour l'orientation du missile et la mise en route du moteur de croisière dans des conditions difficiles de lancement.The gases formed during the ignition of the charge of the pressure generator 11, create an overpressure at the bottom of the container 10 which acts on the end of the rear part of the body 3. The shutter 12, therefore sinks further in the nozzle section 7 by protecting the cruise engine of the missile from the hot gases from the generator 11, which avoids the risk of spontaneous starting of the cruise engine. Part of the gas is ejected through the orifice 16 (cf. FIG. 5) towards the hermetic upper cavity of the container 10. As soon as the pressure under the front cover of the container 10 reaches a critical level, destruction of the front cover occurs. and ejecting debris outward. Once the pressure in the closed space at the bottom of the container reaches the required value, the bolts that hold the missile on the support 14 and the valve 15 of the missile explode, sliding along the surface inner cylindrical guide of the container 10 closes the orifice 16, and the missile shoots upwards and is ejected at the required height (which can reach for example 40m), necessary for the execution of the maneuver for the orientation of the missile and the starting the cruise engine in difficult launch conditions.

Après que le missile a atteint la hauteur requise, ou bien, si cela est possible, sur la partie montante de la trajectoire du missile, on procède à l'exécution des manoeuvres pour l'orientation du missile, c'est à dire le contrôle du tangage, du cap et du roulis. L'exécution de ces manoeuvres est effectuée différemment selon la réalisation des moyens d'orientation du corps annulaire 3.After the missile has reached the required height, or else, if possible, on the rising part of the trajectory of the missile, the maneuvers are carried out for the orientation of the missile, i.e. the control pitch, heading and roll. The execution of these maneuvers is carried out differently depending on the embodiment of the means for orienting the annular body 3.

Pour la première variante du premier mode de réalisation (figure 1, figure 3), après l'allumage par le bloc électronique du missile du générateur de gaz annulaire 4, le jet de gaz chaud arrive simultanément par toutes les conduites de gaz 17, applique les manchons annulaires 23 contre les extrémités de la tringle 18 (les manchons 23, de ce fait, "hermétisent" les jeux du joint amovible), et est éjecté des tuyères 5, en créant des forces réactives, dirigées tangentiellement par rapport au corps annulaire 3, perpendiculairement à son axe, c'est à dire dans un plan perpendiculaire à l'axe du missile 1. La régulation de ces forces de réaction est effectuée simultanément avec la régulation des forces aérodynamiques à l'aide de l'entraînement unique qui commande la rotation des gouvernes 2, liées cinématiquement par les fourchettes en forme de "V" 19 aux tringles 18, qui tournent autour des axes 20. Dans la position neutre des gouvernes 2, qui est représentée sur la figure 1, le gaz arrive dans toutes les tuyères de tous les couples tuyères 5 en quantités égales et la résultante des forces de réaction est égale à zéro (cf. figure 3). En cas de déviation d'une des gouvernes 2 selon un angle maximum (25-30 degrés) d'un côté ou de l'autre, la tringle 18 tourne d'environ 10 degrés, et tout le jet du gaz qui émane de la conduite de gaz 17 n'arrive que dans une des tuyères 5 du couple correspondant. Ainsi, la position angulaire des gouvernes 2 commande la position angulaire de la tringle correspondante 18, et la répartition du jet de gaz entre les tuyères 5 du couple correspondant s'effectue proportionnellement à la position angulaire de la tringle 18, et crée de ce fait des forces de réaction de même signe que dans les plans aérodynamiques de la gouverne 2, assurant la commande du missile en tangage, cap et roulis.For the first variant of the first embodiment (Figure 1, Figure 3), after the ignition by the electronic block of the missile of the annular gas generator 4, the jet of hot gas arrives simultaneously through all the gas pipes 17, applies the annular sleeves 23 against the ends of the rod 18 (the sleeves 23, therefore, "seal" the games of the removable seal), and is ejected from the nozzles 5, by creating reactive forces, directed tangentially with respect to the annular body 3, perpendicular to its axis, that is to say in a plane perpendicular to the axis of the missile 1. The regulation of these reaction forces is carried out simultaneously with the regulation of the aerodynamic forces using the single drive which controls the rotation of the control surfaces 2, kinematically linked by the "V" shaped forks 19 to the rods 18, which rotate around the axes 20. In the neutral position of the control surfaces 2, which is shown s ur in figure 1, the gas arrives in all the nozzles of all the nozzle couples 5 in equal quantities and the resultant of the reaction forces is equal to zero (cf. figure 3). In the event of deviation of one of the control surfaces 2 at a maximum angle (25-30 degrees) on one side or the other, the rod 18 rotates by approximately 10 degrees, and all the jet of gas which emanates from the gas line 17 only arrives in one of the nozzles 5 of the corresponding torque. Thus, the angular position of the control surfaces 2 controls the angular position of the corresponding rod 18, and the distribution of the gas jet between the nozzles 5 of the corresponding torque is carried out in proportion to the angular position of the rod 18, and thereby creates reaction forces of the same sign as in the aerodynamic planes of the control surface 2, ensuring the control of the missile in pitch, heading and roll.

Pour la deuxième variante du premier mode de réalisation du corps annulaire 3 (figures 4, 6 et 7), le principe de création des forces de réaction de direction est analogue à celui qui est mentionné ci-dessus. La différence réside uniquement dans le fait que dans la deuxième variante, la rotation de la tringle 18 est commandée par la rotation de la gouverne 2, ce qui provoque la rotation de la broche 25 (cf. figure 7). La position angulaire de la broche 25 détermine la quantité du gaz qui arrive dans chaque tuyère 5 du couple, et donc la valeur de la résultante des forces de réaction dans le couple de tuyères.For the second variant of the first embodiment of the annular body 3 (FIGS. 4, 6 and 7), the principle of creation of the steering reaction forces is similar to that which is mentioned above. The difference lies only in the fact that in the second variant, the rotation of the rod 18 is controlled by the rotation of the control surface 2, which causes the rotation of the spindle 25 (cf. FIG. 7). The angular position of the spindle 25 determines the quantity of gas which arrives in each nozzle 5 of the couple, and therefore the value of the result of the reaction forces in the couple of nozzles.

Pour le deuxième mode de réalisation du corps annulaire 3 (figure 8), le principe de création des forces de réaction qui commandent le missile 1 est un peu différent de celui qui est décrit ci-dessus. L'orientation du missile 1 est effectuée sans participation des gouvernes aérodynamiques 2, grâce à la mise en route à un instant donné des moteurs à réaction à impulsion, commandés par exemple directement par le calculateur du bloc électronique du missile. Le basculement du missile en tangage et en cap est assuré par la mise en route des moteurs à impulsion les plus puissants des rangées 31-32, dont les tuyères produisent des forces de réaction orientées d'une façon radiale. La direction du plan de basculement du missile est déterminée par les moteurs à impulsion de faible puissance des rangées 29 et 30, dont les tuyères produisent des forces de réaction tangentes au corps annulaire 3.For the second embodiment of the annular body 3 (FIG. 8), the principle of creation of the reaction forces which control the missile 1 is a little different from that which is described above. The orientation of the missile 1 is carried out without the participation of the aerodynamic control surfaces 2, thanks to the start-up at a given instant of the pulse reaction engines, controlled for example directly by the computer of the electronic block of the missile. The rocking of the missile in pitch and heading is ensured by the start-up of the most powerful impulse motors of rows 31-32, whose nozzles produce reaction forces oriented in a radial manner. The direction of the rocking plane of the missile is determined by the low-power pulse motors of rows 29 and 30, whose nozzles produce reaction forces tangent to the annular body 3.

A la fin de la manoeuvre d'orientation du missile en direction de la cible, le moteur de croisière du missile se met en route. Les gaz produits lors du fonctionnement du moteur de croisière éjectent facilement l'obturateur de protection 12 (cf. figures 1, 4 et 8) et après cela, sont éjectés librement par la section de tuyère 7 du corps annulaire 3, augmentant la vitesse du missile. Etant donné que le profil de la section de tuyère 7 est en continuité avec le profil de la tuyère 6 du moteur de croisière, le divergent de la tuyère du moteur de croisière est optimisé, ce qui augmente l'impulsion de la force de réaction du moteur de croisière en fonctionnement et compense une perte éventuelle de vitesse, due à la présence de la masse inerte du corps annulaire 3, représentant les moyens d'orientation, qui a déjà rempli son rôle. Ainsi, le missile emporte la masse inerte suffisamment loin de l'aire de lancement sans consommation énergétique supplémentaire et, si nécessaire peut l'éjecter du missile à un moment donné et en un lieu donné. Pour ce faire, il faut procéder à la destruction des boulons explosifs 8 et, à l'aide des pyro-poussoirs 9 (cf. figure 4) créer une impulsion initiale, nécessaire à l'éjection hors du missile de la masse inerte du corps annulaire 3 comportant les moyens d'orientation qui ont déjà rempli leur rôle, le moteur de croisière étant alors en fonctionnement;At the end of the orientation maneuver of the missile towards the target, the cruise engine of the missile starts. The gases produced during the operation of the cruise engine easily eject the protective shutter 12 (cf. FIGS. 1, 4 and 8) and after that, are freely ejected by the nozzle section 7 of the annular body 3, increasing the speed of the missile. Since the profile of the nozzle section 7 is in continuity with the profile of the nozzle 6 of the cruise engine, the divergence of the nozzle of the cruise engine is optimized, which increases the momentum of the reaction force of the cruise engine in operation and compensates for a possible loss of speed, due to the presence of the inert mass of the annular body 3, representing the orientation means, which has already fulfilled its role. Thus, the missile carries the inert mass far enough from the launch pad without additional energy consumption and, if necessary can eject it from the missile at a given time and in a given place. To do this, it is necessary to destroy the explosive bolts 8 and, using the pyros 9 (see Figure 4) create an initial impulse, necessary for the ejection from the missile of the inert mass of the body annular 3 comprising the orientation means which have already fulfilled their role, the cruising engine then being in operation;

En conclusion, la présente invention permet, avec un minimum de consommation d'énergie, l'interception d'une cible apparue subitement à proximité de l'aire de lancement, située dans un environnement difficile, et en même temps de réduire à un minimum l'incidence néfaste du lancement du missile sur l'aire de lancement en éliminant la nécessité de l'éjection de la masse inerte des moyens d'orientation après exécution de leur fonction. L'invention peut être appliquée aussi bien à des missiles de grandes dimensions que de faibles dimensions. En outre, l'invention permet, moyennant une modification minimale des missiles existants à lancement incliné, de leur conférer toutes les qualités mentionnées ci-dessus. Les trois modifications proposées dans les cas particuliers de réalisation du système de commande de lancement et d'orientation du missile, sont, du point de vue des paramètres qualitatifs, équivalentes. Le choix de l'une ou l'autre est déterminé par la spécificité du missile qui devra les utiliser. Les moyens utilisés dans des circonstances données peuvent être moins appropriés dans d'autres conditions.In conclusion, the present invention allows, with a minimum of energy consumption, the interception of a target appeared suddenly near the launch pad, located in a difficult environment, and at the same time reduce to a minimum the harmful impact of the missile launch on the launching area by eliminating the need to eject the inertial mass from the orientation means after the performance of their function. The invention can be applied to both large and small missiles. In addition, the invention allows, with minimal modification of existing missiles with inclined launch, to give them all the qualities mentioned above. The three modifications proposed in the particular cases of implementation of the launch and orientation control system of the missile, are, from the point of view of the qualitative parameters, equivalent. The choice of one or the other is determined by the specificity of the missile which will have to use them. The means used in given circumstances may be less appropriate in other conditions.

Claims (12)

  1. System for launching and for steering flying vehicles, comprising launching means and a flying vehicle, the said vehicle including aerodynamic control surfaces (2) with their drive and steering means located in the rear part of the flying vehicle, and including at least one gas generator (4) and nozzles (5) which are connected to it, characterized in that it includes an annular body (3) which is rigidly connected to the body of the flying vehicle (1), the steering means being located in the annular body, the internal surface of the annular body having a truncated-cone shape and being coated with a thermally insulating material, forming a section of nozzle whose profile is continuous with the profile of the nozzle of the flying vehicle's sustainer motor.
  2. System according to Claim 1, characterized in that the annular body includes means (8, 9) ensuring that it is ejected by the flying vehicle in flight.
  3. System according to Claim 1 or 2, characterized in that the jet thrust nozzles of the steering means are located in the same plane, perpendicular to the longitudinal axis of the section of nozzle.
  4. System according to one of Claims 1, 2 and 3, characterized in that the launching means are produced in the form of launching containers (10) with front and rear covers, the internal volume of which is cylindrical and is intended to accommodate the flying vehicle, the pressure generator (11) being located at the bottom of the container, which is closed by the rear cover (13) and a protective shroud (12) having a frustoconical lateral surface, the profile of which follows at least certain parts of the surface of the section of nozzle of the annular body, the rear part of the annular body including a peripheral closure member (15) whose outside diameter is equal to the inside diameter of the container, the container including a support to which weakened elements are fixed, these weakened elements being intended for fixing the annular body above the pressure generator.
  5. System according to Claim 4, characterized in that the protective shroud has a convex shape, its convex part being turned towards the sustainer motor.
  6. System according to one of Claims 4 and 5, characterized in that the launching container includes, in the region for fixing the annular body, an ejection orifice (16) whose dimensions are chosen so as to take into account the gas flow passing through the clearance which is formed around the closure member of the annular body, the front cover of the container being produced so as to be fragmented under a given pressure which develops inside the container.
  7. System according to one of Claims 3, 4, 5 and 6, characterized in that it is provided with rods (18) fixed to the annular body, the gas generator (4) being annular and connected to the nozzles of the steering means via gas ducts (17) formed in the annular body, the nozzles (5) all being identical, grouped in pairs in the same plane, the nozzles of each pair being oriented in opposite directions and mechanically connected to one end of the corresponding rod, ensuring that the gas jet from the common gas duct in the annular body is distributed among the nozzles, each rod being connected by its other end to a corresponding control surface (2), thereby making joint rotation possible.
  8. System according to Claim 7, characterized in that it is provided with annular sleeves (23) made of heat-resistant material which are located close to the output end of the corresponding gas duct (17), it being possible for these sleeves to move longitudinally inside this end, each rod being fixed to the annular body in its central part via its rotation spindle (20), each pair of nozzles being produced in the form of sharply-bent ducts, with frustoconical output ends and inlet orifices facing the outlet of the common gas duct, and the diameters of which are equal to the inside diameter of the heat-resistant annular sleeves, the contact surfaces of the first end of each rod and of the annular body being thermally insulated.
  9. System according to Claim 5, characterized in that each pair of nozzles is produced in the annular body in the form of a single straight channel having frustoconical end-pieces, the annular body including radial orifices (24), the axis of which, on one side, passes through the centre of the corresponding straight channel, is perpendicular to the axis of the latter and lies in the same plane and, on the other side, is perpendicular to the axis of the output pipe of the corresponding common gas duct and lies in a second plane, and, finally, the axis of these orifices lies on the intersection of the two planes, each rod being fixed to the annular body at one of its ends by means of a pin (25), which is coated with a heat-stable composite material, and arranged so as to ensure rotation in the corresponding radial orifice, which is coated with a heat-insulating layer, the layer of composite material on each pin having an ejection orifice (27A) in order to ensure distribution of the gas jet between the nozzles in the pair.
  10. System according to one of Claims 3, 4, 5 and 6, characterized in that the steering means are produced in the form of jet thrusters (29k to 32k) located in the annular body in rows (29 to 32) which are uniformly distributed heightwise, each thruster nozzle being oriented perpendicularly to the longitudinal axis of the gas duct of the annular body, each row being formed by the thrusters of a single type and having the same size.
  11. System according to Claim 10, characterized in that at least the lower-power thrusters form one row, the end-pieces at the outlet of the nozzles of these thrusters being directed tangentially with respect to the annular body.
  12. System according to Claim 10, characterized in that, in a first row of thrusters, the end-pieces at the outlet of the nozzles of these thrusters are all directed tangentially in a given direction with respect to the annular body and in that, in another row, including thrusters of the same type as those in the first row, the end-pieces at the outlet are all directed in the opposite direction to that of the end-pieces in the first row.
EP95936617A 1994-10-27 1995-10-27 Missile launching and steering system Expired - Lifetime EP0737297B1 (en)

Applications Claiming Priority (5)

Application Number Priority Date Filing Date Title
RU94040077 1994-10-27
RU94040077 1994-10-27
RU95110350A RU2082946C1 (en) 1995-07-03 1995-07-03 Missile take-off and orientation actuating system
RU95110350 1995-07-03
PCT/FR1995/001423 WO1996013694A1 (en) 1994-10-27 1995-10-27 Missile launching and steering system

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EP0737297A1 EP0737297A1 (en) 1996-10-16
EP0737297B1 true EP0737297B1 (en) 1997-10-08

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JP (1) JP3692537B2 (en)
KR (1) KR100404037B1 (en)
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DE (1) DE69500842T2 (en)
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NO (1) NO310637B1 (en)
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KR100404037B1 (en) 2004-03-24
NO962653D0 (en) 1996-06-21
JP3692537B2 (en) 2005-09-07
FI111032B (en) 2003-05-15
DK0737297T3 (en) 1997-12-22
ES2107921T3 (en) 1997-12-01
FI962638A (en) 1996-08-23
IL115749A0 (en) 1996-01-19
FI962638A0 (en) 1996-06-26
AU708097B2 (en) 1999-07-29
NO962653L (en) 1996-08-27
EP0737297A1 (en) 1996-10-16
WO1996013694A1 (en) 1996-05-09
UA27153C2 (en) 2000-02-28
TW319825B (en) 1997-11-11
DE69500842D1 (en) 1997-11-13
JPH09507567A (en) 1997-07-29
DE69500842T2 (en) 1998-02-26
US5823469A (en) 1998-10-20
IL115749A (en) 2000-02-29
AU3848195A (en) 1996-05-23
NO310637B1 (en) 2001-07-30

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