RU2235282C1 - Rocket missile - Google Patents

Abstract

FIELD: defense industry; production of rocket missiles.

SUBSTANCE: the invention is dealt with the field of arms and may find application in the guided-missile systems of a short-range radius of action. A rocket missile contains a separated engine with a sustainer stage, the rear part of which is joined with a located in the engine sleeve, in which there is an installed ignition device. Between the sustainer stage rear butt, walls of the sleeve and the ignition device there is a cavity connected with a combustion chamber of the separated engine through a throttle valve. The rear part of the sustainer stage is supplied with the heat-insulating pan linked with the sleeve with the help of a destructible element located along the axis of the rocket missile. Production of a rocket missile in compliance with the offered invention will allow to ensure reduction of the temperature dispersion of the maximum speed of the rocket missile due to application of a self-adjustable mode of operation of the offered design.

EFFECT: the invention will allow to ensure reduction of the temperature dispersion at the maximum speed of the rocket missile.

3 cl, 1 dwg

Description

The proposed invention relates to the field of armaments and may find application in short-range missile systems.

Known rocket shell [1], adopted by the authors as an analogue, with a detachable engine, containing a device for docking different stages of the shell. It consists of a transition conical fairing with a central tube covering the march stage of the projectile, which allows the engine to slide parallel to the axis in the opposite direction of the march stage, and a separation mechanism in the form of a slotted channel device on the conical part of the fairing, communicating on one side with the cavity of the march stage, and on the other, with a stream of air rushing onto the shell.

The analogue design is most acceptable in unguided rockets of multiple launch rocket systems, since a large cantilever departure of the marching stage with a small embedment depth into the engine, which leads to disturbances during separation, practically does not affect the accuracy of hitting the areal target, and is not acceptable for anti-aircraft guided missiles shells with engines made of composite material, since disturbances arising from the separation of the march stage from the engine can lead to a sharp change in the angle at aka, which leads to the exit of the marching stage from the beam of the control system, which leads to the loss of the projectile. In addition, with a large cantilever departure of the marching stage relative to the engine, it is necessary to have a solid connection of different-caliber rocket stages, the length of the marching step significantly affects the longitudinal stability of the projectile in flight. The operation of the separation mechanism to a large extent depends on the velocity of the projectile at the time of separation and on the angle of attack of the projectile. The minimum separation time is achieved at a zero angle of attack, an increase in the angle of attack leads to an increase in the non-uniformity of the velocity field at the entrance to the slots, the appearance of asymmetric lateral forces, which leads to an increase in friction in the docking unit at the time of separation and an increase in disturbances during separation.

The closest analogue adopted by the authors for the prototype of the invention is a rocket projectile [2] containing a marching stage with a cowl and a motor docked with it using a separation mechanism. A thin-walled metal cup is installed in the engine, reinforced from the outside with a heat-shielding material, in the bottom of which an igniter is installed. The back of the marching stage is connected to the glass and partially recessed in the engine.

The design of the prototype allows firing at long ranges without using an additional marching engine due to the separation of the engine with a caliber larger than the marching stage caliber due to the fact that when the starting engine is separated, the midship and side surfaces of the rocket are reduced and, as a result, drag and friction. Deepening part of the sustainer stage into the engine increases the longitudinal stability of the projectile in flight.

Due to the diversity of the marching stage and the starting engine, their separation occurs due to the difference in aerodynamic forces acting on the marching stage and the starting engine after the engine thrust level in the recession area becomes less than the frontal aerodynamic drag force.

However, when using a projectile in a wide range of ambient temperatures, the spread in the maximum velocity of the projectile reported to it by the starting engine increases significantly. The presence of this scatter is caused primarily by a change in the total thrust impulse and the total operating time of the detachable starting engine when the initial charge temperature changes (when the temperature decreases, the impulse decreases, and the total operating time increases), as well as by a change in air density with a change in the initial temperature. A decrease in the total momentum with an increase in the operating time leads to a decrease in the maximum velocity of the projectile due to a decrease in traction and an increase in the impulse of the drag force, which, together with an increase in air density, leads to the fact that at negative ambient temperatures the maximum velocity of the projectile is much lower than with positive . The presence of a large temperature spread of the maximum velocity of the projectile most negatively affects when firing at long ranges at a retreating target. At the same time, when the engine is separated due to aerodynamic force, there is a significant dispersion in separation time caused by the presence of control torque on the rudders of the march stage, which also leads to an increase in the dispersion of the maximum projectile speed.

Thus, the objective of the invention is to reduce the temperature spread of the maximum velocity of the projectile, which will improve efficiency when firing at long ranges at a retreating target in a wide temperature range.

The task is achieved in that in a rocket with a detachable engine containing a marching stage, the rear of which is docked with a glass placed in the engine, in which an igniter is installed, in contrast to the prototype, a cavity is made between the rear end of the marching stage, the walls of the glass and the igniter communicating with the combustion chamber of the detachable engine by means of a throttle channel, while the back of the marching stage is equipped with a pallet connected to the glass p zrushaemym element, the cross-sectional area which is defined as

where σ is the tensile strength of the material of the destructible element in tension or in shear;

R _floor - the pressure in the cavity between the rear end of the march stage, the walls of the glass and the igniter;

- площадь торца маршевой ступени;

- the area of the end of the march stage;

m _D is the mass of the engine;

n _x - the magnitude of the axial overload acting on the projectile at the time of destruction of the element;

- сила сопротивления, действующая на двигатель;

- resistance force acting on the engine;

- давление в камере сгорания двигателя;

- pressure in the combustion chamber of the engine;

- площадь критического сечения сопла двигателя;

- the area of the critical section of the engine nozzle;

C _m - calculated value of the thrust coefficient of the engine nozzle, defined as

where k is the ratio of the specific heat of the combustion products used in the engine fuel (an integral characteristic of each fuel is given in the documentation);

P _a , P _n - pressure of the products of fuel combustion at the exit section of the nozzle and the ambient pressure, respectively;

F _a - the area of the output section of the nozzle of the projectile engine;

μ is the calculated value of the flow rate of the engine nozzle, defined as

,

,

- для радиусного входа, r₂ - радиус входной части сопла, r_кр - радиус критического сечения сопла.where ζ _x = (θ _in / 90) ^0.924 - for a nozzle with a conical inlet, θ _in - the angle of inclination of the generatrix of the input cone of the nozzle to the longitudinal axis of the projectile;

- for a radial inlet, r ₂ is the radius of the inlet part of the nozzle, r _kr is the radius of the critical section of the nozzle.

In this case, the pallet is thermally insulated, and the destructible element is located along the axis of the projectile.

The implementation of a missile in accordance with the alleged invention will allow:

- to reduce the temperature spread of the maximum velocity of the projectile due to the self-regulating mode of operation of the proposed design.

In the process of dispersal of the projectile at the site of the engine’s operation, the marching stage is fastened with a glass with a destructible element. Act on him:

- tensile force from the pressure of the combustion products of the fuel coming from the combustion chamber of the working engine through the throttle channel into the cavity, made between the rear end of the sustainer stage, the walls of the glass and the ignition device;

- tensile inertial force acting on the engine side in flight;

- the tensile force of aerodynamic drag acting on the engine, or the friction force when moving along the container;

- compressive axial component of the engine traction force.

In the process of pressure accumulation in the cavity, the tensile force increases in proportion to the increase in pressure in the cavity between the rear end of the sustainer stage, the walls of the cup and the igniter. The tensile inertial force and the tensile force of aerodynamic drag increase in proportion to the increase in axial overload and the velocity of the projectile. The compressive axial component of the engine traction force is determined by the nature of the change in traction over time and in most cases is almost constant.

At the end of the thrust and pressure drop in the combustion chamber of the engine at the end of its operation, due to the reduction of axial overload, the tensile force acting on the destructible element decreases, the tensile force of aerodynamic resistance remains almost unchanged. Since the pressure in the cavity between the rear end of the marching stage, the walls of the glass and the igniter device decreases more slowly than the engine thrust, the tensile force increases by the time of separation.

Separation of the march stage and the engine occurs when the condition

,

,

where F _bit. - force destruction of the element;

m _d is the mass of the engine;

n _x - axial overload acting on the projectile at the time of separation:

,

,

where GΣ is the current value of the total mass of the projectile;

- сила сопротивления, действующая на двигатель;

- resistance force acting on the engine;

R is the engine thrust;

- сила давления пороховых газов в полости между задним торцом маршевой ступени, стенками стакана и воспламенительным устройством на торец маршевой ступени.

- the pressure force of the powder gases in the cavity between the rear end of the sustainer stage, the walls of the glass and the ignition device on the end of the sustainer stage.

After the destruction of the element, separation is carried out by expanding the gas accumulated into the cavity under the action of the resistance force acting on the starting engine and the friction force when the march step moves along the glass.

The area of the throttle channel, the volume of the cavity and the material of the destructible element are selected from the condition that the tensile force on the destructible element of compressive forces is not exceeded during operation of the engine

and also from the condition of obtaining the minimum residual impulse of the engine after the separation of the stages at the minimum air temperature

,

,

where m _d - the mass of the empty engine;

V _{department. D} - engine compartment speed;

τ _sec. - engine separation time;

J _rst. - the residual impulse of engine thrust after its separation from the sustainer stage, which eliminates the possibility of an empty engine impact after its separation at the rear end of the sustainer stage.

The cross-sectional area of the destructible element is assigned to the engine operating conditions at the extreme value of the negative temperature of the specified range. Since with an increase in the initial temperature, the pressure pulse in the combustion chamber of the engine increases, then the pressure in the cavity between the rear end of the sustainer stage, the walls of the glass, and the ignition device increases proportionally to it. Since the nature of the pressure drop in the combustion chamber of the engine, its thrust and the overload acting on the projectile depends on the initial temperature less than the pressure level in the combustion chamber, the cavity between the rear end of the march stage, the walls of the cup and the igniter, as well as the level of thrust and overload, then with increasing temperature the moment of the onset of separation shifts in time towards higher values of thrust, i.e. the separation process with an increase in the initial temperature begins earlier. This leads to the fact that the engine is separated with a partially unburnt charge, which leads to a decrease in the total thrust impulse, the maximum velocity of the projectile and a decrease in the temperature spread of the maximum velocity of the projectile;

- reduce the momentum of the drag force at negative air temperatures by reducing the separation time and thereby increase the final velocity of the projectile at the launch site and thereby also reduce the temperature spread of the maximum velocity of the projectile;

- to ensure reliable separation of the sustainer stage with the engine, regardless of the ratio of their calibers and reduce the time of separation due to the pressure force of the powder gases taken from the combustion chamber of the engine into the cavity made between the rear end of the sustainer stage, the walls of the cup and the igniter;

- to provide due to the application of a thermal insulation coating to the pallet, protection of the rear end of the sustainer stage from heating during the whole time of the engine operation and thereby prevent its expansion, which can lead to an increase in friction force or jamming of the sustainer stage when the engine is separated from it;

- eliminate the skew of the pallet due to the location of the destructible element along the axis of the projectile;

- to ensure the autonomy of the separation unit.

The invention is illustrated by the scheme of the missile shown in the drawing.

The proposed missile contains a marching stage 1 and a detachable engine 2. The rear end 3 of the marching stage 1 is docked with the glass 4 installed in the engine 2, in which the igniter device 5 is installed. Between the rear end 3 of the marching stage 1, the walls of the glass 4 and the ignition device 5 is made a cavity communicating with the combustion chamber of the engine by means of a throttle channel 8. At the rear of the marching stage, a pallet 6 with a heat-insulating coating 9 is fixed, connected to the glass with a destructible element 7, located along the axis of the projectile.

The work of the proposed design is as follows. After engine 2 is triggered, it accelerates the projectile to a predetermined maximum speed. During engine operation, through the throttle channel 8, the products of fuel combustion enter the cavity made between the rear end 3 of the march stage 1, the walls of the glass 4 and the ignition device 5. As the charge burns out in the engine, the pressure in the cavity increases, axial overload acting on the engine grows 2, and aerodynamic drag. When this increases the tensile force acting on the destructible element 7. From destruction it protects the compressive force from the thrust of the engine. At the end of engine operation, the thrust and axial overload drop sharply, and the pressure in the cavity between the rear end 3 of the march stage 1, the walls of the cup 4 and the igniter 5 changes slightly. As a result, the element 7 is destroyed and the engine 2 is separated from the sustainer stage 1 due to the expansion of the gas accumulated into the cavity under the action of the aerodynamic drag force acting on the engine and the friction force when the sustainer stage moves along the glass.

The volume of the cavity, the area of the throttle channel and the cross-sectional area of the element to be destroyed are selected in each case by calculation and are specified in the process of mining.

Thus, in the proposed technical solution, the temperature dispersion of the maximum velocity of the projectile is reduced, which allows to increase efficiency when firing at long ranges at a retreating target in a wide temperature range. Reducing the temperature spread of the maximum speed can reach ~ 10 ÷ 15%. Ultimately, the reliability of the rocket and the effectiveness of the complex.

Sources of information

1. Application of France No. 2629583, MKI F 42 B 15/00, published 06.10.89, the analogue.

2. Patent RU No. 2133444, MKI F 42 B 15/10, published July 20, 1999, bulletin No. 20 - prototype.

Claims (3)

1. A projectile with a detachable engine, comprising a marching stage, the rear of which is docked with a glass placed in the engine, in which an igniter is installed, characterized in that a cavity communicating with the combustion chamber is made between the rear end of the marching stage, the walls of the glass and the ignition device detachable engine by means of a throttle channel, while the rear part of the sustainer stage is equipped with a pallet connected to the glass by a destructible element, the area of the transverse with whose cross section is determined from the relation

, where σ is the tensile strength of the material of the destructible element in tension or in shear; R _floor - the pressure in the cavity between the rear end of the march stage, the walls of the glass and the igniter;

- the area of the end of the march stage; m _D is the mass of the engine; n _x - the magnitude of the axial overload acting on the projectile at the time of destruction of the element;

- resistance force acting on the engine;

- pressure in the combustion chamber of the engine;

- the area of the critical section of the engine nozzle; C _m is the calculated value of the thrust coefficient of the engine nozzle; μ is the calculated value of the flow rate of the engine nozzle. 2. The missile according to claim 1, characterized in that the tray is made insulated. 3. A missile according to claim 1 or 2, characterized in that the destructible element is located along the axis of the projectile.