RU2814225C2 - Interceptor missile - Google Patents

Abstract

FIELD: weapons.

SUBSTANCE: invention relates to armament, namely to interceptor missile. Interceptor jet comprises a multistage solid-propellant rocket engine consisting of a plurality of working chambers arranged along a common axis, docked to each other due to the fact that the upper part of the combustion chamber of the previous stage is tightly adjacent to the inner surface of the supercritical part of the nozzle of the next stage. At the edge of the supercritical part of the nozzle of the working chamber of each stage there are at least three cutouts, they are distributed along the nozzle circumference and covered with steering gates arranged on the outer side of the nozzle shell, which are fixed on the nozzle shell by means of a hinge located in the upper part of the damper and providing the possibility of controlled deflection of the damper to the outside. There is also a drive for deflection of said flaps by means of motion transmission mechanisms acting synchronously on flaps of all stages located on the common vertical line and not preventing free undocking of the stages. Steering flaps of all rocket stages of the multistage engine, located on one vertical line, are connected to each other by a chain consisting of levers performing the function of limiters of the value of possible opening of the flaps, and wire rods connecting said levers to each other, as well as to servo electric drive located on upper stage. Links of said chain belonging to adjacent rocket stages are connected by means of two levers connected to each other by means of a fork with an open slot oriented longitudinally relative to the rocket axis and located on arms of said levers moving transversely. Plane of rotation of the levers performing the function of the shutter opening limiter is located in relation to the shutter surface at an angle greater than the friction angle. Steering flaps located at least on one of the verticals are attached to the nozzle surface by means of a biaxial ball hinge, that is with possibility of deflection by angle not only in plane passing through rocket axis, but also in the plane transverse to the rocket axis, and they are connected to two independent servo electric drives acting on the corresponding two edges of one shutter.

EFFECT: possibility to control thrust vector direction during the whole rocket flight.

1 cl, 6 dwg

Description

The invention relates to rocket science and can be used to create aerospace defense (ASD) systems based on small-sized hypersonic kinetic-action guided missiles.

In modern conditions, when the high accumulated energy of nuclear offensive weapons becomes incompatible with the possibility of their use, the preferred means of protecting states in conditions of political confrontation is the development of defense systems that combine low impact energy with high precision and efficiency of its control. This also applies to means of intercepting air and space targets.

The most universal means of intercepting air and space targets are guided missiles equipped with rocket engines, because they can operate in a wide range of altitudes - from zero to interplanetary, as well as speeds reaching 10 km/sec or more.

However, the shortcomings of existing interception missile systems are associated with the large launch weight and high cost of production of interceptor missiles, which does not allow aerospace defense assets to be distributed throughout the country with sufficient density and, in particular, to bring them closer to all vital infrastructure facilities of the country. Interception of space objects over the border of a state’s territory is not suitable because space is open to all countries. It is necessary to identify all space targets. But identifying hostile targets does not mean the need to immediately intercept them. It is necessary to wait for the manifestation of aggressiveness, which can only be revealed when flying over the territory and at a close distance from strategically important infrastructure facilities.

In order to have time to intercept a target that has shown aggressive intentions, for example, one that has begun to descend or has become separated, it is necessary to place the interceptor launch sites as close as possible to all protected objects (no more than a few km). Moreover, the number of interceptors for each object must be no less than the expected number of warheads (and possibly decoys) into which the enemy object is divided. It is clear that when using large, multi-tonnage missiles, this is impossible to do, because There will not be enough state resources to manufacture and maintain large and expensive missiles.

The large tonnage of existing interceptor missiles is due to two reasons:

1) the need to achieve high - hypersonic and even cosmic speeds for interception, which requires multi-stage;

2) the need to have sufficient energy from the warhead to ensure hitting the target with inevitably limited guidance accuracy, which involves the use of a warhead whose radius exceeds the miss value.

From these conditions, the value of the final mass, the required number of stages, as well as the ratio of the starting mass to the final mass are determined. For example, if the area of the critical vulnerable part of the target is 1 square meter, then the required density of distribution of ready-made destructive elements (GPE) after their dispersion should be 1 element per square meter. If the probable miss of the guidance system is 10 m, then 300 GGE will be required. With a GGE mass of 10 g, the required warhead mass will be 3 kg. If the rocket's final stage payload ratio is 50%, then the final rocket mass will be 6 kg. To intercept space targets, the minimum acceleration speed of the warhead must be at least 9 m/sec, which approximately corresponds to the launch of the payload into low Earth orbit (LEO). To obtain such a final speed with a single-stage rocket at a speed of gas outflow from the nozzle of 3 km/sec, it will be necessary (according to the Tsiolkovsky formula) a launch mass of the rocket 6 * 20 = 120 kg (since the number “e” = 2.7 to the power equal to the ratio final velocity to the exhaust velocity, i.e. 9/3, equals 20). However, this is ideal when the entire service mass is included in the final mass of the rocket structure, which is 3 kg. If the rocket is single-stage, then the final mass of the structure must include the mass of the liquid-propellant rocket engine fuel tanks or the mass of the working chamber of the solid propellant rocket engine, which can accommodate the entire volume of fuel. Moreover, the interceptor must have a high starting acceleration - at least 10 Gs - in order to pick up speed quickly enough. Those. the engine thrust must be more than 10 times the launch weight of the rocket. This makes it impossible to achieve escape velocity with a single stage. Those. the rocket must be multi-stage. In practice, they make two or three steps. However, this significantly makes the starting mass heavier than the above ideal one. Moreover, three engines, three stage launch systems, three control systems and two undocking systems are required. And all this for the sake of a warhead weighing 3 kg. Economic optimization forces the mass of the warhead to be increased several times. But even at the same time, the cost of accelerating a unit of mass to 9 km/sec will remain very high. Those. the task of creating a fairly cheap and effective aerospace defense system with the launch positions of interceptor missiles as close as possible to all important elements of the country’s infrastructure using rockets with traditional rocket engines liquid propellant rocket engines or solid propellant rocket engines cannot be solved with any possible amount of funding. But even if such a system is built, its use, in the event of a massive attack, will lead to destruction commensurate with the damage from the means of attack against which the aerospace defense system is used.

However, a control system for a rocket with a multi-stage solid propellant engine is known (see RU 2021113927 A, publ. 11/17/2022).

The purpose of the invention is to eliminate the design disadvantage of an interceptor missile built on the basis of a multi-stage solid propellant rocket engine. Namely, the task is to ensure the ability to economically control the direction of the thrust vector throughout the entire flight, i.e. ensuring the possibility of intensive maneuvering with insignificant expenditure of working fluid for maneuvering throughout the entire flight, including the most critical final stage of maneuvering the interceptor in an airless environment. At the same time, the task is also set to miniaturize the interceptor projectile as much as possible while maintaining the possibility of hitting a target, for example, by causing damage that reduces the effectiveness of its warhead.

An interceptor rocket is proposed, containing a multi-stage solid fuel rocket engine, consisting of a plurality of working chambers located along a common axis, coupled to each other due to the fact that the upper part of the combustion chamber of the previous stage fits tightly to the inner surface of the supercritical part of the nozzle of the next stage. The purpose of the invention is achieved by the fact that at the edge of the supercritical part of the nozzle of the working chamber of each stage there are at least three cutouts, distributed along the circumference of the nozzle and covered, each, by steering flaps located on the outside of the nozzle shell, each fixed to the nozzle shell using a hinge, located in the upper part of the damper and providing the possibility of controlled deflection of the damper to the outside. There is also a drive for deflecting said steering valves through a motion transmission mechanism that acts synchronously on the valves of all stages located on a common vertical and does not interfere with the free uncoupling of the stages. In this case, these flaps perform the function of rudders, the action of which is based on the fact that the pressure force of the gas flow of the supercritical part of the nozzle on the indicated flaps covering the cut out supercritical part of the nozzle, when the flaps are deflected, decreases with increasing deflection of the steering flap. This is equivalent to the appearance of a transverse force of the opposite sign acting on the bottom of the rocket and causing the angular acceleration of the rocket in pitch or yaw. The maximum value of this force is equal to the pressure force of the gas flow on the cut out part of the area of the supercritical shell of the nozzle.

In a particular variant of the design, the steering valves of all rocket stages of a multistage engine, located on the same vertical, are connected to each other by a chain consisting of levers that act as limiters on the amount of possible opening of the valves, and wire rods connecting these levers to each other, as well as to a servo-electric drive located on the top step. In this case, these levers interact with the steering valves at an angle that provides two-way transmission of movement, from the lever to the valve and from the valve to the lever. Moreover, the connection of the links of the said chain belonging to adjacent rocket stages is made by means of two levers connected to each other by means of a fork with an open groove oriented (relative to the rocket axis) longitudinally and located on the shoulders of the said levers moving transversely. This connection provides a combination of the possibility of transmitting vertical motion between the wire rods of adjacent stages with the possibility of unhindered separation of the stages vertically during the separation of the spent stage.

In this way, the most difficult control problem is solved - controlling the orientation of the rocket in terms of heading and pitch angles.

To solve the problem of roll control due to the propulsion engine, in a particular version of the design, the steering dampers located on one of the verticals are attached to the nozzle surface by means of a biaxial or ball (universal) joint, i.e. with the possibility of angular deflection not only in the plane passing through the axis of the rocket, but also in the plane transverse to the axis of the rocket. In this case, an additional servo drive is introduced to control the angle of deflection of the steering valve in this plane. The occurrence of a force tangential to the rocket axis is explained by the fact that the vector of the total gas pressure on the valve is practically perpendicular to the surface of the valve. Thus, at least one of the dampers will create an alternating force located tangential to the rocket axis, which is sufficient to control the roll, i.e. to stabilize the rocket in rotation around its axis.

In a particular embodiment, the specified additional drive for controlling the angular position of the steering damper in the transverse plane is an additional mechanism similar to that described above, located on the outer surface of the nozzle, but interacting with the second edge of the same steering damper, secured by a biaxial hinge. Thus, all steering valves located on one of the vertical chains will be synchronously controlled at two deflection angles. In this case, the unidirectional component of the deflection of the edges of the specified damper will work to control the heading and pitch, and the differential component will work to control the roll.

Coordinated control of the specified angles of at least three steering flaps distributed around the circumference makes it possible to set the roll, pitch and yaw angles required for control, i.e. provides complete control necessary to stabilize and maneuver the missile. At the same time, during the control process, the levers located on the overlying stages also make the above movements, synchronously with the dampers of the lower, currently operating stage. But they are outside the flow of gases and do not experience force effects (except for small forces from the air during the atmospheric flight phase). Thus, the force required for control is not summed up in stages, and it is possible to control the direction of the engine thrust vector, as well as the orientation of the missile along the roll angle throughout the flight, including the extra-atmospheric section, where intensive maneuvering is required to intercept fast-flying targets. In this case, maneuvering is carried out by the main engine. Moreover, fuel consumption increases only to the extent of the loss of thrust created by the indicated nozzle cutouts, which have a relatively small area.

The invention is illustrated by the following detailed description of an example of a structural embodiment and six figures.

Figure 1 shows a schematic cross-section of the proposed interceptor missile.

Figure 2 shows the appearance of the proposed projectile from the side, where only one of the three vertical lines of the location of the shutters along the steps is visible.

Figure 3 shows an enlarged image of the levers and their connection to each other and to the servo-electric drive.

Figure 4 shows a section through the transverse plane A-A shown in Figure 1. The location of the steering valves and the levers and wire rods that control them along the circumference is shown.

Figure 5 illustrates the effect of the lever on the possible angle of deflection of the steering valve. The dependence of the possible angle of deflection of the steering valve, under the influence of gas pressure, on the position of the lever roller is shown.

Figure 6 explains the principle of adding three control forces: “a; b; c" to obtain two vectors of shear force Χ; Y, which determine the pitch and heading moments.

The rocket engine with the proposed control system contains a plurality of working chambers 1, docked together by means of a tight fit of the outer surface 2 of the combustion chamber of the previous stage with the inner surface of the supercritical part 3 of the nozzle of the next stage. This creates a stack of rocket stages, the length and composition of which can be quickly changed depending on the individual flight mission. This is ensured by the unification of the dimensions and shape of the joining zones of the working chambers of all stages, as well as the presence of a prepared set (store) of stages with different parameters for thrust and combustion time. In the fire bottom of all stages, except the last one, there is a mechanism 4 for transferring combustion to the overlying stage, initiated at the final stage of fuel burnout of the previous stage. This can be either a breakthrough membrane, exposed when the layer of solid fuel becomes thinner, or a valve for bypassing hot gases into the overlying combustion chamber, opened by an automatic device based on a signal from a rocket overload reduction sensor or a signal from an off-board source, received, for example, via a radio channel.

At the edge of the shell of the supercritical part of the nozzle of each stage there are at least three cutouts 5 (see Figs. 1 and 3) distributed around the circumference of the nozzle exit. These cutouts are covered by steering flaps 6, attached to the outer surface of the nozzle by means of hinges 7, 8 and 9 (see Fig. 4). Moreover, hinges 7 and 8 are cylindrical (uniaxial), i.e. each have only one axis of rotation, which ensures only one degree of movement of the damper. And the hinge 9 is biaxial, providing the ability to deflect the damper at two angles, namely in a plane passing through the longitudinal axis of the rocket, as well as in the transverse plane.

Deflections of the dampers are limited by mechanisms consisting of a system of levers and wire rods connecting the dampers with four servo-electric drives 10 located on the upper rocket stage. The action of these mechanisms is to control the limiters on the amount of possible opening of the steering dampers 6. In this case, two-way transmission of movement must be ensured - both from the limiter to the damper and from the damper to the limiter. This is ensured by the choice of geometric parameters of the mechanism for connecting the limiter with the damper. Moreover, the damper mechanisms located on the same vertical act synchronously.

In a particular design variant, the mechanism, in a section that is intermediate between adjacent stages, contains a three-arm lever 11 (see Figs. 2 and 3) installed on the lower edge of the working chamber of the rocket stage, and a two-arm lever 12 installed on the body of the working chamber of the underlying stage near the specified three-arm lever 11. The condition for the two-way transmission of movements between the three-arm lever 11 and the steering damper 6 is ensured by the fact that the angle “a” between the plane of rotation of the lever 11 and the tangent to the surface of the steering damper is selected greater than the angle of friction of the lever 11 with the surface of the damper (see. Fig.5). Moreover, these three-arm and two-arm levers of adjacent stages are connected to each other by means of a fork 13, the open groove 14 of which is oriented in the longitudinal (relative to the axis of the rocket) direction. The specified two-arm 12 and three-arm 11 levers of adjacent steps are connected by wire rods 15 with similar mechanisms of the underlying and overlying steps, respectively. The rod 16 of the last stage is connected to the servo-electric drive 10 (see Figs. 2 and 3). This device ensures two-way transmission of the tension force of the wire rods along the entire vertical chain of steps. And at the same time, the possibility of unhindered separation of stages in flight is ensured.

The shoulder 17 of each three-arm lever 11 serves as a limiter for the deflection of the steering valve and is equipped at the end with a roller 18, which serves to reduce friction (see Fig. 3). Moreover, the lower stage is also equipped with a three-arm lever, because its third arm with a fork can be used when increasing the number of stages, for example, if it is necessary to intercept a higher-speed and higher-altitude target.

The steering dampers 6, attached by uniaxial hinges 7 and 8 (Fig. 4), have only one lever 11 each, which limits the angle of deflection of the damper in the transverse plane. And the steering damper with a biaxial hinge 9 is equipped with two levers 11, which corresponds to two degrees of mobility of this damper (see Fig. 4). In this case, the synchronous component of the deflection of the two levers 11 of such a damper controls the course of the possible deflection of the damper in the longitudinal plane passing through the axis of the rocket, and the differential component controls the deflection of the damper in the plane transverse to the axis of the rocket. This deviation corresponds to the appearance of a shoulder of gas pressure on the damper, i.e. the appearance of a moment that creates rotation of the rocket around its axis. This is used to control the missile's roll angle.

Servo-electric drives 10 are located in a depression on the surface of the working chamber of the upper stage. To reduce aerodynamic drag, the levers are positioned, if possible, parallel to the surface of the working chambers. Additionally, all mechanisms can be covered with a casing (casing not shown).

In addition to the main engine, in the head of the rocket there is a rocket control unit 19, as well as receivers for control and positioning signals of the rocket, or an autonomous navigation and control system. There is also a warhead, which, in a particular case, can be a set of ready-made striking elements (GPE) 20 located under the fairing, equipped with a powder charge 21, which serves to dissipate the GGE and is triggered at a given distance to the target, determined by an on-board proximity sensor with the target or external flight tracking system.

The proposed interceptor projectile functions as follows. By initiating the ignition of the lower stage, they launch from a tubular guide (mortar launch). In this case, the control system immediately comes into operation, the action of which is to stabilize or change the orientation of the longitudinal axis of the rocket in the angles of course, pitch and roll by changing the deflection angles of the lower stage flaps 6 along all available degrees of freedom, limited by hinges 7, 8 and 9. The indicated deflections of the dampers 6 are produced by longitudinal movement of four chains of wire rods using servo-electric drives 10. In this case, the servo-electric drive 10 only affects the limitation of the angle of possible deflection of each steering damper 6, and the deflection of the steering damper itself is carried out under the influence of pressure Ρ of gases in the supercritical part of the nozzle (see Fig.5). The gas pressure on the dampers 6 occurs both due to the underexpansion of gases in the supercritical part of the nozzle, and due to the curvature of the wall of the supercritical part of the nozzle, causing a decrease in the divergence cone of the gas flow. Moreover, when the valve 6 is deflected, the gas pressure on it decreases, down to zero, because the damper in this case reduces the degree of deviation of the flow of gases emerging from the corresponding cutout 5 of the nozzle. A decrease in pressure on valve 6 is equivalent to the appearance of a transverse force in the opposite direction, i.e. is equivalent to the deviation of the thrust vector of the operating stage of a solid propellant engine.

In the case under consideration, the use of four control channels to control three moments - heading, pitch and roll, has one degree of redundancy. This redundancy manifests itself in the form of the ability to obtain the required components X and Υ of the vector (see Fig. 6) with different lengths of the three vectors “a”, “b” and “c” located at an angle of 120 degrees. However, this redundancy is necessary because flaps 6, unlike conventional rudders, cannot create an alternating transverse (to the rocket axis) force.

The fourth chain of rods is necessary to create a controlled tangential force used for roll control, i.e. to control the rotation of the rocket around its axis. Tangential force, unlike transverse radial force, is sign-alternating. This allows you to control the rotation of the rocket around its axis due to the propulsion engines, which makes it possible to abandon the use of any additional low-thrust engines or flywheel torque orientation engines.

When the fuel of the lower solid propellant stage approaches complete burnout, the thrust decreases, and this can serve as a signal for the activation of the combustion transfer mechanism to the upper stage. For example, when the layer of solid fuel becomes thinner, the membrane of mechanism 4 may break. In this case, the flame breaks into the higher combustion chamber and produces ignition. Pressure appears in the overlying chamber, which pushes out the underlying spent stage. In this case, the two-arm and three-arm levers of adjacent stages, connected to each other by a fork 13 with an open groove 14, are freely disengaged, and the operation of the control system is automatically transferred to the overlying stage without interruption, because separation of the underlying stage occurs at the moment when there is already pressure in the working chamber of the upper stage, due to which the separation of the stages occurs.

The task of intercepting a target is made easier due to the increased radius of dispersion of the GGE 20, which can be controlled by changing the distance to the target when the dispersive charge 20 is detonated. At the same time, due to the high hypersonic speeds of the relative movement of the projectile and the target, the GGE work like a kinetic weapon. In this case, the kinetic energy of GPE is many times greater than the energy of chemical bonds of any material. The kinetic energy of the GGE, due to its high speed, does not have time to dissipate in the material over a large distance from the channel axis and acts similarly to the action of a cumulative ammunition, which can penetrate through any armor and any warhead, disrupting the functioning of the warhead. In particular, in this case, an abnormal activation of a nuclear warhead charge is possible, leading to an asymmetrical implosion with a decrease in the force of the explosion. In addition, the effectiveness of the attack is reduced due to the warhead not reaching the target.

Thus, the proposed projectile is not inferior in efficiency and damage radius to large multi-stage missiles. The simplicity of the design, as well as the reduction in the size of the rocket, will reduce the cost of creating a solid one, i.e. distributed throughout the territory, aerospace defense, also designed for a massive attack, complicated by a large number of decoys. In addition, a small-sized, and therefore cheap to mass-produce, interceptor missile allows for an increased number of misses, as is the case in artillery, where shells are significantly cheaper than multi-stage missiles.

Claims (1)

An interceptor missile containing a multi-stage solid fuel rocket engine, consisting of a plurality of working chambers located along a common axis, interconnected due to the fact that the upper part of the combustion chamber of the previous stage fits tightly to the inner surface of the supercritical part of the nozzle of the next stage, while On the edge of the supercritical part of the nozzle of the working chamber of each stage there are at least three cutouts distributed along the circumference of the nozzle and each covered by steering flaps located on the outer side of the nozzle shell, fixed to the nozzle shell using a hinge located in the upper part of the damper and providing the possibility of controlled deflection of the damper in outer side, and there is also a drive for deflecting said dampers by means of motion transmission mechanisms acting synchronously on the dampers of all stages located on a common vertical and not interfering with the free uncoupling of the stages, characterized in that the steering dampers of all rocket stages of a multistage engine, located on the same vertical, are connected between each other by a chain consisting of levers that serve as limiters on the amount of possible opening of the valves, and wire rods connecting these levers to each other, as well as with a servo-electric drive located on the upper stage, and the connection of the links of this chain belonging to adjacent rocket stages is made by means of two levers connected to each other by means of a fork with an open groove oriented longitudinally relative to the rocket axis and located on the shoulders of said levers moving transversely, and the plane of rotation of the levers acting as a valve opening limiter is located relative to the valve surface at an angle greater than the friction angle, and located on at least one of the verticals, the steering flaps are attached to the surface of the nozzle by means of a biaxial ball joint, i.e. with the possibility of angular deflection not only in a plane passing through the axis of the rocket, but also in a plane transverse to the axis of the rocket, and they are connected to two independent servo-electric drives acting on the corresponding two edges of one damper.

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