RU2422663C1 - Solid-propellant rocket engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: solid-propellant rocket engine includes chamber with nozzle, front bottom, solid-propellant charge from bundle of cartridges, which is attached to the front bottom, priming composition in perforated metal tube, which is attached to the bottom, and nozzle closure. In perforated metal tube on the side of the front bottom there is perforated section throughout the length of not more than the half of the tube length, after which a pin with diameter of 0.1 to 0.3 of inner diameter of the tube is installed perpendicular to tube axis. Perforation holes are covered with cover from gas-permeable material, which is fixed on outer surface of the tube. Shell with pyrotechnic tracer agent arranged inside it, which protrudes beyond the end of cartridge bundle, is installed on the tube on the nozzle side. Nozzle closure is made from flexible material in the form of membrane with disc-shaped flanging, central boss with a hole and annular groove. Flanging of the closure is fixed in the connector made in supercritical part of the nozzle, and its groove is made on the side of the nozzle bell.

EFFECT: invention allows increasing energy characteristics of rocket engine and safety of its use.

3 cl, 2 dwg

Description

The invention relates to the field of armament, in particular to rocket engines (RD) of solid fuel for mobile systems intended for firing from the shoulder, for example grenade launcher or flamethrower samples.

Such complexes are subject to increased requirements to ensure the parameters of the “habitability” of the shooter, in terms of the gas-flame flow when the projectile leaves the launcher, and in limiting the parameters of the temperature and acoustic effects on the shooter. These requirements are especially stringent when shooting at half-closed positions and in rooms of limited volume. These requirements are met through various technical solutions, for example, such as special designs of charges and ignition devices, as well as the design of assemblies flying out of the RD nozzle during firing, such as nozzle plugs, and other devices that determine the parameters of factors affecting the operator when firing .

In addition, when starting rocket engines are used for melee systems, a characteristic feature is the limited conditions for the implementation of the energy characteristics of solid fuels. The resulting energy losses are due to the operation of the RD in an unsteady pulsed mode, in which propellant fuel particles fly out of the nozzle unburned.

Launch rocket engines (SRDs) used in grenade launchers, flamethrowers and ATGMs mainly use a “brush” charge design in which a beam of solid fuel elements (checkers) is fastened on one side to the bottom of the SRD. This design is the most promising and widely used for this class of ammunition.

Known technical solutions used in the SDR for shooting “from the shoulder” are the use of solid fuel charges of the “brush” design in the form of a bundle of tubular elements fastened to the bottom of the SDR (French patents No. 2181178, F42C, 1974, No. 2439174, F42B, 1996, Russian patent No. 2211354, F02K, 2002).

A technical solution is known (RF patent No. 2062428), which provides for the ignition of SDR charges by using a prechamber perforated metal tube installed in the central charge channel with an igniter composition inside. In this design, the use of an igniter device with a prechamber tube with uniformly located perforations in the area of the channel of a multi-plate solid fuel charge leads to intense radial loads on the surface of the charge elements, which reduces the strength characteristics of the structure and can lead to abnormal charge operation under extreme loads.

Known rocket engine (RF patent No. 2251628 - prototype), containing a "brush" charge, bonded to the front bottom, and an axial igniter device in the Central channel of the charge, consisting of a metal prechamber perforated tube in the form of two sections separated by a partition with a throttle hole. This design allows you to stabilize the process of ignition of the charge and to ensure the weakening of the radial effects of ignition products on the charge elements of solid fuel. The disadvantage of this technical solution is to increase the dimensions and mass of the prechamber tube, as well as an increase in heat loss during the passage of combustion products through an additional second section. As a result of this, an additional increase in the mass of the igniter composition and an increase in the dimensions of the igniter device are required, which leads to a decrease in the useful volume of the chamber occupied by solid fuel, and, consequently, to a decrease in energy with the same dimensions of the taxiway.

The problem solved by the claimed invention is to increase the efficiency of the rocket engine and ensure the safety of its use.

The technical result consists in increasing the energy characteristics of the rocket engine, and reducing the parameters of negative ergonomic factors affecting the firing.

The technical result is achieved in that the solid fuel rocket engine comprises a chamber with a nozzle, a front bottom, a charge of solid fuel from a bunch of checkers attached to the bottom, an igniter in a perforated metal tube fixed to the bottom, and a nozzle plug. A perforated section is made in the perforated metal tube from the front bottom side at a length of not more than half the length of the tube, behind which a pin with a diameter from 0.1 to 0.3 of the inner diameter of the tube is installed perpendicular to the axis of the tube. The perforation holes are closed by a shell of a gas-permeable material (for example, a cotton cloth) fixed on the outer surface of the tube, and a glass with a pyrotechnic tracer placed inside the protrusion protruding beyond the end of the block of checkers is mounted on the tube from the nozzle side. The nozzle cap is made of an elastic material (e.g., rubber) in the form of a membrane with a plate flanging, a central boss with a hole and an annular groove, the flanging being pinched in the connector made in the supercritical part of the nozzle, and the groove is made on the nozzle socket side.

In the tube, at the base of the cup with a tracer located inside, longitudinal grooves with a length of 0.1 to 0.3 of the inner diameter of the tube can be made.

The igniter composition may consist of a fine-grained fraction located in the area of the perforated tube in the area from the front bottom to the pin, and a coarse-grained fraction located in the tissue shell in the area from the pin to the tracer.

Perforation from the front bottom of a metal tube with an igniter composition on the length of not more than half the length of the tube allows you to increase the gas intake rate of the combustion products of the igniter composition in the area of the solid fuel tubes at the base of the charge, while also increasing the time of exposure of the combustion products to the tubular charge elements due to their flow from the closed zone of the tube to the zone with the perforation section, which in turn contributes to an increase in the efficiency of the process of burning a solid charge th fuel.

The introduction of a pin installed perpendicular to the tube with the igniter composition, partially overlapping the tube cross-section, makes it possible to control the ignition composition moving speed together with the combustion products, which ensures a reduction in the number of unburned particles of the ignition composition, especially for the coarse-grained fraction.

The installation of an easily destructible elastic nozzle plug in the supercritical part of the nozzle, with a weakened cross section in the form of an annular groove, ensures opening of the nozzle with the lowest possible pressure, which helps to reduce the shock wave parameters of the supersonic gas flow in the nozzle and positively affects the required habitability parameters (acoustic pressure and temperature exposure).

The installation of a pyrotechnic tracer in the pre-nozzle volume increases the energy characteristics of the rocket engine due to the afterburning of unburned particles of solid fuel passing through the nozzle and the impact of the high-temperature combustion products of the pyrotechnic composition on the gas stream.

The implementation of the longitudinal grooves of the tube in the perforated tube at the base of the glass additionally stabilizes the ignition process due to the action of combustion products of the igniter flowing from the longitudinal grooves on the end sections of the solid fuel charge tubes from the nozzle side.

Figure 1 shows a rocket engine in the context.

Figure 2 shows the igniter device - a perforated tube with an igniter composition.

The rocket engine shown in Fig. 1 consists of a chamber 1: a nozzle socket 2, a front bottom 3, to which a solid fuel charge 4, consisting of a bunch of checkers, is attached. In the central part of the bottom 3 is fixed a metal perforated tube 5, which is an ignition device. At the end of the tube 5, from the nozzle 2 side, a glass 6 is installed, in which the pyrotechnic tracer 7 is placed. Sealing of the rocket engine is provided by an elastic membrane 8, which is clamped by its flanging between the chamber 1 and the nozzle socket 2.

The igniter device shown in figure 2, contains a metal perforated tube 5 with holes located in the area from the front bottom 3, closed by a shell of gas-permeable material 9, mounted on the outer surface of the tube 5. Inside the tube 5 is an igniter composition 10 and the initiating device 11, for example an electric igniter. The igniter composition 10 is divided into parts by a pin 12 fixed inside the tube 5. On the pin 12 inside the tube 5, an electric igniter 11 is fixed in the perforation zone. A possible embodiment of the igniter composition 10 may be its separation (by pin 10) into parts containing various fractions, including fine-grained and coarse-grained. On the tube 5 from the side of the glass 6, longitudinal grooves can be made.

The work of the proposed rocket engine of solid fuel is as follows:

When applying an electric pulse to the electric igniter 11, the igniter composition 10 is ignited, which burns the shell of the gas-permeable material 9, the combustion products begin to flow out through the openings of the tube 5, igniting the solid fuel charge checkers 4.

When the specified pressure is reached, the elastic membrane 8 is opened over a weakened cross-section (annular groove), and the process of expiration of the combustion products of charge 4 from the nozzle socket 2 begins.

During the ignition of composition 10, the ignition front propagates inside the perforated tube 5 towards the nozzle portion of the taxiway. After opening the membrane 8, the products of combustion of the igniter composition 10 from the closed zone inside the tube 5 flow into the perforation zone and provide fuel for the ignition and combustion of the charge of solid fuel 4. Some of the products of combustion of the igniter composition 10 expire from the grooves in the tube 5, providing more intense ignition of the end sections charge 4 and ignition of the pyrotechnic tracer 7.

The time-stretched process of deformation and opening of the elastic membrane 8 and the extended process of ignition of the charge 4 together contribute to a decrease in the intensity of the initial “pop” at the start of the taxiway, thereby reducing the level of acoustic pressure on the firing point.

At the initial moment of ignition of the charge 4, the pyrotechnic composition of the tracer 7 is ignited. The high-temperature combustion products generated during the operation of the tracer 7 provide pyrotechnic support for the combustion process and gas outflow from the RD nozzle, contributing to the afterburning of the unburnt particles of charge 4, thereby increasing the completeness of combustion and energy RD.

The presented design of the taxiway is widely tested as part of grenade launchers by full-scale tests in various conditions, including firing by the operator from the shoulder, with positive results.

Claims (3)

1. A rocket engine of solid fuel containing a chamber with a nozzle, a front bottom, a charge of solid fuel from a bunch of checkers attached to the front bottom, an igniter composition in a perforated metal tube fixed to the bottom, and a nozzle cap, characterized in that in a perforated metal tube from the front bottom side, a perforated section is made at a length of not more than half the length of the tube, behind which a pin is installed perpendicular to the axis of the tube with a diameter of 0.1 to 0.3 of the inner diameter of the tube, The perforations are covered by a sheath of gas-permeable material fixed on the outer surface of the tube, and on the tube on the nozzle side there is a glass with a pyrotechnic tracer inside, protruding beyond the end of the block of checkers, the nozzle cap is made of elastic material in the form of a membrane with a plate flanging, a central boss with hole and an annular groove, and the flanging is pinched in the connector made in the supercritical part of the nozzle, and the groove is made from the side of the nozzle socket. 2. The rocket engine according to claim 1, characterized in that in the tube at the base of the cup longitudinal grooves are made with a length of 0.1 to 0.3 of the inner diameter of the tube. 3. The rocket engine according to claim 1 or 2, characterized in that the igniter composition consists of a fine-grained fraction with an electric igniter located in the area of the perforated tube in the area from the front bottom to the pin, and a coarse-grained fraction placed in the tissue shell in the area from the pin to tracer.

Images