RU2482321C1 - Solid-propellant rocket engine start system and pressure intake of solid-propellant rocket engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: solid-propellant rocket engine start system includes pyro cartridges installed in the housing cover plate, augmentor tube, igniter and its attachment assembly. The igniter attachment assembly includes a shell, a ring and longitudinal screws passing through the ring and shell walls and screwed into the threaded seats on the cover plate. Open end face of the shell contacts the igniter, and the ring presses the igniter flange to the open shell end face. The augmenter tube passes through the axial hole at the shell bottom, which is equal to its diameter. Pressure intake of solid-propellant rocket engine is made around the augmenter tube of the start system and includes seats of the telemetric measurement system, which are made in the cover plate of the rocket engine housing and have the channels gas-connected to inner cavity of solid-propellant rocket engine housing, and the screen covering these channels and forming the bottom of the start system shell. On the side of outer surface of the shell bottom there are annular perforated projections dividing the volume between the shell and the heat-insulating coating into several coaxial headers. Slots are made in the heat insulating coating between the igniter attachment screws.

EFFECT: inventions allow reducing the weight and dimensions of the start system and pressure intake, improving reliability and simplifying the manufacturing procedure.

8 cl, 5 dwg

Description

The invention relates to rocket technology and can be used to create a rocket engine of solid fuel (solid propellant rocket engine).

It is known that the solid propellant rocket engine must contain a launch system [Design of solid propellant rocket engines / Ed. ed. LN Lavrova - M .: Engineering, 1993 - 215 p., Ill., Page 165, Fig. 4.1] and, in many cases, a pressure intake [ibid, page 189, Fig. 5.2].

The launch system is designed to ignite the charge of solid fuel on the command supplied to the squibs, and contains a squib (one or more) installed in the solid propellant rocket housing (usually in the front cover), an afterburner, an ignitor, and an igniter mount.

The pressure intake is designed to select intracameral pressure to the sensors of the telemetry measurement system (STI) and reduce the thermal effect on the sensors to an acceptable level in the process

- fire test rigs of solid propellant rocket engines (at the stage of testing solid propellant rocket engines and during periodic tests at the stage of mass production),

- flight tests (at the stage of testing the aircraft (LA) with solid propellant rocket engine),

- regular launches of the aircraft (in case information on the in-chamber pressure is required by the aircraft control system).

The pressure inlet contains STI nests located in the solid-propellant housing (cover) with channels connected to the internal cavity of the solid-propellant housing, and a screen covering these channels from direct exposure to combustion products. In accordance with the above tasks (taking into account the need for periodic testing of any randomly selected from a manufactured batch of solid propellant rocket engines), it is advisable to perform a pressure intake on all manufactured solid rocket engines (even if information on the in-chamber pressure of the aircraft control system is not required). In this case, the STI sockets on the solid-state solid propellant motors in the standard version are closed with plugs, and in the case of periodic testing, the plugs are replaced with STI sensors.

The RDTT launch system and the RDTT pressure intake require the execution of numerous bosses in the solid-state solid-propellant body (for example, on the front cover), increasing the mass of the structure and complicating the manufacturing technology of the solid-state solid propellant.

The combination of the solid propellant launch system and the solid propellant pressure intake into a single unit is known in the solid propellant pump housing [RF Patent 2230926]. However, this circuit is based on the use of a non-combustible igniter body as an STI screen. The fireproof igniter housing, as a rule, is used only on small-sized solid propellant rocket engines, has a large mass. The scheme under consideration does not allow the use of an igniter with a combustible housing, which is widely used in large-size solid propellant rocket motors.

The closest in technical essence and the achieved positive effect to the proposed invention is a solid propellant rocket launcher system and a solid rocket motor pressure trap [RF patent No. 2424442]. The solid propellant rocket launcher system contains a squib (squib) mounted in the cover of the solid propellant rocket motor, afterburner, igniter and its mount. The RTTT pressure inlet, made around the afterburner of the RDTT launch system, contains the telemetry system jacks with channels connected to the internal cavity of the RDTT body located in the cover of the RDTT body and a screen covering the channels. The disadvantage of this launch system and pressure intake is that this design is designed for an igniter having a mount in the form of a rigid threaded ring. An example of such igniters is presented in [Design of solid propellant rocket engines / Ed. ed. LN Lavrova - M.: Mechanical Engineering, 1993 - 215 p., Ill., Pages 168, 169, Fig. 4.5, 4.6]. The attachment point of a number of igniters does not have a rigid threaded ring, but is made in the form of a flange [Designs of solid propellant rocket engines / Under total. ed. LN Lavrova - M .: Mechanical Engineering, 1993 - 215 p., Ill., Page 168, Fig. 4.4]. The fastening of such igniters requires a different circuit of the solid propellant rocket launcher. Accordingly, it is rational to integrate the design of the pressure intake with the modified scheme of the launch system.

An object of the present invention is to reduce the mass and dimensions of the design of the solid propellant rocket ignition system with an ignitor equipped with a flange, and the solid rocket pressure transducer, solid propellant housing cover, simplifying their manufacturing technology, and increasing reliability.

The essence of the invention “RDTT launch system” is that in the RDTT launch system containing pyrocartridges (squib) installed in the cover of the solid propellant rocket engine, afterburner, igniter and its mount, the mount contains a cup, the open end of which is in contact with the ignitor, a ring pressing the igniter by means of a flange made on the igniter to the open end of the cup, and longitudinal screws passing through the ring and the walls of the cup and screwed into the threaded sockets on the lid. The afterburner passes through an axial hole equal in diameter to the bottom of the glass. The glass may be provided with a shoulder fixing the ignitor flange in the radial direction, while the ring is provided with an annular selection in which the shoulder is placed. Lugs surrounded by a heat-shielding coating can be connected to the lid of the solid-propellant rocket motor through an annular bridge, and threaded sockets are made in lugs. Openings can be made in the glass, communicating the internal cavity of the glass with the internal cavity of the solid propellant rocket motor. On the glass and the heat-resistant coating of the lid, segmented selections between the screws can be made.

The essence of the invention, "the RTTT pressure intake" is that in the RTTT pressure intake made around the afterburner of the RDTT launch system, which contains the telemetry system sockets in the cover of the RDTT body with channels connected to the internal cavity of the RDTT body, a screen covering the channels, a screen covering the channels forms the bottom of the cup of the solid propellant rocket launcher. On the side of the outer surface of the bottom of the glass, annular perforated protrusions are made on it, dividing the volume between the glass and the heat-protective coating into several coaxial collectors. In the heat-shielding coating, grooves are made between the screws of the igniter mounting. A seal can be installed between the afterburner of the solid propellant rocket engine and the forming surface of the axial hole in the bottom of the beaker. The perforation of the annular protrusions can form holes and grooves arranged in such a way that opposite to each hole and groove is an unperforated portion of the annular protrusion.

The technical result in the solid propellant rocket launcher is achieved by the fact that the igniter attachment unit comprises a bottom having a cup, into the internal cavity of which an afterburner is brought out. Thus, the bottom of the cup forms two cavities isolated from each other - the inner cavity of the cup, inside which the force of the flame from the afterburner is transferred to the igniter when the solid propellant is started, and the cavity between the bottom and the heat-shielding coating, which makes it possible to use it to form a pressure intake integrated with launch system. The screws that press the igniter with the ring through the flange made on the igniter to the open end of the cup fasten the igniter, the cup and the pressure intake elements at the same time. A bead made on the glass secures the igniter flange in the radial direction. Ring sampling in the ring provides radial fixation of the ring relative to the glass. The threaded sockets into which the screws are screwed are made in bosses connected to the lid by means of an annular jumper. This during operation of the solid propellant rocket motor provides a thermal barrier between the screw heads exposed to the combustion products and the cover. A simplification of the manufacture of the lid is also provided: in one piece with the lid (an annular jumper connected to it), an axisymmetric ring is sharpened, from which the bosses are then milled. The lugs are surrounded by a heat-protective coating of the lid (like the entire lid). The implementation in the glass of radial or inclined holes communicating the internal cavity of the glass with the internal cavity of the solid propellant rocket motor provides:

- the ability to check the tightness of the internal cavity of the solid propellant rocket chamber by boosting this cavity (both through the socket of the squib and through the socket of the telemetry measurement system);

- the possibility of drying the internal cavity of the solid propellant rocket motor through the pyro cartridge socket;

- reduction of pressure casting in the internal channel of the afterburner and in the internal cavity of the glass during the operation of the squib.

Performing segmented samples between the screws on the glass and the heat-insulating coating reduces the weight of the structure.

The proposed design of the solid propellant rocket launcher system minimizes the transverse dimensions of its main elements and, as a result, reduces their mass, simplifies manufacturing technology, and improves the reliability of the structure. Prerequisites for reducing the transverse dimensions of the pressure intake are provided.

The technical result in the RTTT pressure intake is achieved by increasing the density of the arrangement of the pressure intake elements made around the afterburner of the RDTT launch system, due to the fact that the screen covering the channels forms the bottom of the nozzle of the RDTT launch system. Moreover, from the side of the outer surface of the bottom of the glass, annular perforated protrusions are made on it, dividing the volume between the glass and the heat-shielding coating into several coaxial (i.e. densely located) collectors. The external collector is in communication with the internal volume of the solid propellant rocket motor housing by means of the fact that grooves are made between the screws of the igniter in the heat-shielding coating. The bosses in the cover of the solid propellant rocket housing for the STI nests are practically at the same minimized radius as the bosses for the pyro cartridge nests. The maximum realization of the potential advantages of combining the launch system and the pressure intake into a single unit is achieved by the fact that the cover of the solid-propellant rocket motor housing is made in the form of a simple axisymmetric (i.e., technological) thin-walled (i.e., having a minimum mass) membrane having only one central boss with minimum radius. In a single central boss, all the necessary sockets are made (for squibs and for STI sensors). The perforation of the annular protrusions is formed by holes and grooves located “in a checkerboard pattern”, i.e. so that opposite to each hole and groove is a non-perforated portion of the annular protrusion. This reduces the thermal effect on the STI sensors, i.e. improves the reliability of STI. The specified configuration of the holes and grooves provides a uniform distribution of pressure around the circumference of the coaxial collectors during sudden changes (or fluctuations) in the pressure inside the chamber, i.e. lack of circulation of combustion products, and the performance of the STI when slagging part of the longitudinal blind holes, which increases the reliability of the STI. A seal is installed between the afterburner of the RDTT launch system and the generatrix of the axial hole in the bottom of the beaker. This ensures the tightness of the bottom and eliminates the influence of a pressure jump in the afterburner and the inner cavity of the glass on the readings and performance of the STI.

The proposed design of the pressure intake RTTT minimizes the transverse and axial dimensions of its main elements and, as a result, reduces their mass, simplifies manufacturing techniques, and improves the reliability of the structure.

This technical solution is not known from the patent and technical literature.

The invention is illustrated by the following graphic material:

figure 1 shows a view from the outside of the cover of the solid propellant rocket motor with elements of the solid propellant rocket launcher system and solid propellant pressure intake placed on it;

figure 2 shows a longitudinal section along aa of figure 1;

figure 3 shows a longitudinal section along BB of figure 1;

figure 4 shows a longitudinal section along bb In figure 1;

figure 5 shows a cross section along G-D of figure 2.

The solid propellant rocket engine (RDTT) launch system comprises one or more pyro-cartridges 1 mounted in boss sockets 2, made (see FIG. 2) in the cover 3 of the solid propellant rocket motor housing. Sockets of the squibs 1 are made from the side of the outer surface of the cover 3. From the side of the internal cavity of the solid propellant body on the cover 3 there is an afterburner 4 connected to the sockets of the squibs 1. The afterburner 4 is directed to the igniter 5, which is fastened by the attachment unit 7 made on the ignitor 5 of the flange 6 . The attachment unit 7 comprises a cup 8 having a bottom 9. The open end 10 of the cup 8 contacts the igniter 5. The igniter 5 is pressed against the open end 10 of the cup 8 by means of the ring 11. Through the ring 11 and the walls of the cup 8, longitudinal screws pass 12, screwed into the threaded sockets 13 on the cover 3. The threaded sockets 13 are made in bosses 14 connected to the cover 3 by means of an annular bridge 15 and are surrounded by a heat-shielding coating 16. The afterburner 4 passes through an axial hole 17 equal in diameter to the bottom 9 of the glass 8. The glass 8 is equipped with a shoulder 18, fixing the flange 6 of the igniter 5 in the radial direction. The ring 11 is equipped with an annular sample 19, in which the shoulder 18 is placed. In the glass 8 there are made radial or inclined holes 20 (see Fig. 4), communicating the internal cavity of the glass 8 with the internal cavity of the solid propellant rocket motor. On the glass 8 and the heat-insulating coating 16, segmented selections 21 were made between the screws 12.

The pressure intake of the solid propellant rocket engine (RDTT) is made around the afterburner 4 of the RDTT launch system. It contains the boss 2 of the cover 3 of the solid-state solid-state drive housing of the socket 22 of the telemetry measurement system with channels 23 connected to the internal cavity of the solid-state solid-state housing (see FIG. 3). The screen covering the channels 23 forms the bottom 9 of the glass 8 of the solid-state propellant rocket launcher. From the side of the outer surface of the bottom 9 of the cup 8, on the bottom 9, annular perforated protrusions 24 are made, dividing the volume between the cup 8 and the heat-shielding coating 16 into several coaxial collectors 25. In the heat-shielding coating 16, grooves 26 are made between the screws 12 of the fastener for the igniter 5, which communicate with the collectors 25 s the internal cavity of the solid propellant rocket motor (see figure 4). Between themselves, the coaxial collectors 25 are gas-coupled by perforation of the annular protrusions 24. The perforations of the annular protrusions 24 are formed by holes 27 and grooves 28, which are arranged so that opposite to each hole 27 and groove 28 there is an unperforated portion 29 of the annular protrusion 24 (see Fig. 5). Channels 23 go into the internal collector 25, communicating with the sockets 22 of the telemetry measurement system. Plugs 30 or STI sensors are installed in slots 22. A seal 31 is installed between the afterburner 4 of the RDTT launch system and the generatrix of the axial hole 17 in the bottom 9 of the cup 8.

The device operates as follows. During manufacture (assembly) and ground operation, the solid propellant rocket motor is checked for leaks by pressurizing the internal cavity of the solid rocket motor with test pressure in one of two ways.

According to the first embodiment, the test pressure is supplied through the socket from the igniter 1, the internal channel of the afterburner 4, the openings 20. Then, the pressure passes through the slots 26, manifolds 25, openings 27, slots 28 and channels 23 to the slots 22. The sockets 22 are closed either with plugs 30 or STI sensors.

In the second embodiment, the test pressure is supplied through one of the sockets 22 through channels 23, manifolds 25, grooves 28, openings 27 and grooves 26 into the internal cavity of the solid propellant rocket motor. Then the pressure through the holes 20, the internal channel of the afterburner 4 falls to the squib 1.

Thus, when applying test pressure both through the socket from the squib 1, and through the socket 22, a reliable check of the tightness of all joints of the solid propellant rocket is ensured due to the access of the test pressure to both the squib 1 and the sockets 22 of the STI.

If necessary, during the manufacture (assembly) and ground operation of solid propellant rocket motors, the internal cavity of the solid propellant rocket chamber is dried by repeatedly filling this cavity with dry gas. Filling is done according to one of the two previously described options (i.e., both through the nest from the squib 1, and through the nest 22).

When conducting periodic tests of any accidentally selected from a manufactured batch of solid propellant rocket tubes, plugs 30 (in the event that information on the chamber pressure is not required by the aircraft control system, plugs 30 are installed on the standard solid propellant rocket) are replaced by STI sensors. An engine equipped with either plugs 30 or STI sensors is ready for operation.

During operation of the solid propellant rocket motor (i.e., until its launch), the loads applied to the igniter 5 are absorbed by the ring 11 and screws 12 of the attachment unit 7, which fastens the igniter 5 on the cover 3 of the solid propellant rocket motor housing. This ensures structural integrity.

When the command to start the engine, a current pulse is applied to the squib 1 mounted on the solid propellant rocket motor. The force of the flame from the igniter 1 through the afterburner 4 ignites the pyrotechnic composition of the igniter 5. In this case, the pressure build-up in the inner cavity of the nozzle 8 is moderate, which ensures a selected value of the volume of the inner cavity of the nozzle 8 and the hole 20, providing an additional consumption of combustion products from the igniter 1, which reduces maximum pressure peak in the specified cavity. This reduces the load on the mount 7 of the igniter 5. In the process of operation of the ignitor 5, the solid propellant is started.

With further operation of the solid propellant solid propellant rocket igniter 5 burns almost instantly, exposing the inner surface of the glass 8.

The measurement of the parameters of the operation of the solid propellant rocket motor (pressure, pressure pulsation) is ensured by unhindered access of the measured pressure to 22 STI sensors located in the sockets. Moreover, the system, consisting of annular perforated protrusions 24, dividing the volume between the cup 8 and the heat-shielding coating 16 into several coaxial collectors 25, provides a significant reduction in the thermal effect on the STI sensors.

The technical and economic efficiency of the present invention, compared with the prototype, which is selected as a solid propellant rocket launcher and a solid rocket intake [RF patent No. 2424442], consists in reducing the mass and dimensions of the solid rocket launcher with an igniter equipped with a flange, and a solid rocket , RTTT case covers, simplifying their manufacturing technology, improving reliability.

Claims (8)

1. The rocket engine starting engine for solid fuel (RDTT), containing a squib (pyro cartridge) installed in the cover of the solid propellant rocket housing, afterburner, igniter and its mount, characterized in that the mount contains a glass, the open end of which is in contact with the ignitor, a ring pressing the igniter by means of a flange made on the igniter to the open end of the cup, and longitudinal screws passing through the ring and walls of the cup and screwed into the threaded sockets on the lid, the afterburner The cask passes through an axial hole equal in diameter to the bottom of the glass. 2. The RDTT launch system according to claim 1, characterized in that the cup is provided with a shoulder fixing the ignitor flange in the radial direction, and the ring is equipped with an annular sample in which the shoulder is placed. 3. The RDTT start-up system according to claim 1, characterized in that the bosses are connected to the cover of the solid propellant by means of an annular bridge, surrounded by a heat-protective coating, and threaded sockets are made in the bosses. 4. The RDTT start-up system according to claim 1, characterized in that the holes are made in the glass, communicating the internal cavity of the glass with the internal cavity of the solid-state rocket engine. 5. The RDTT start-up system according to claim 1, characterized in that segment selections between the screws are made on the glass and the heat-resistant coating of the lid. 6. The pressure intake of the solid propellant rocket engine (RDTT), made around the afterburner of the RDTT launch system, containing in the cover of the RDTT case the sockets of the telemetry measurement system with channels connected to the internal cavity of the RDTT case, a screen covering the channels, characterized in that the screen covering the channels forms the bottom of the glass of the solid propellant rocket launcher, and from the side of the outer surface of the bottom of the glass, annular perforated protrusions are made on it, dividing the volume between the glass and are heat-protective coating at several coaxial collectors, with a thermal barrier coating between the igniter fixing screws are made grooves. 7. The RTTT pressure intake according to claim 6, characterized in that a seal is installed between the afterburner of the RDTT launch system and the generatrix of the axial hole surface in the bottom of the beaker. 8. The RTTT pressure inlet according to claim 6, characterized in that the perforation of the annular protrusions is formed by holes and grooves arranged so that opposite to each hole and groove is an unperforated portion of the annular protrusion.

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