RU2424442C1 - 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 rocket engine housing, igniter and augmentor tube with igniter attachment thread. Igniter is installed on thread of augmentor tube by means of the head having a connection pipe which joins the above thread. The sleeve some part of inner channel of which has the diameter exceeding outer diameter of connection pipe is installed outside augmentor tube. The other invention of the group refers to pressure intake of solid-propellant rocket engine, which includes holes for telemetric measurement system, which are made in the housing of rocket engine and have the channels gas-connected to inner cavity of rocket engine housing, and the screen covering these channels. Rocket engine housing and sleeve edge are equipped with a step providing the gap between rocket engine housing and sleeve. Screen is made in the form of hollow cylinder installed coaxially to sleeve and covering the gap. Gap together with screen forms the annular header to which channels of telemetric measurement system are led. ^ EFFECT: inventions allow improving the reliability of start system and pressure intake of solid-propellant rocket engine, simplifying their manufacturing technique and reducing weight and overall dimensions. ^ 9 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. L.N. 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 telemetric measurement system (STI) and reduce the thermal effect on the sensors to an acceptable level in the process:

- fire bench tests 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 launch system and a pressure intake made on the lid [Design of solid propellant rocket engines / Under total. ed. L.N. Lavrova. - M.: Mechanical Engineering, 1993. - 215 s, ill., Page 124, Fig. 2.55, diagram at the top left]. The combination in the prototype of the launch system and the pressure intake into a single unit is designed to reduce the weight and dimensions of the cover. The disadvantages of these devices are associated with the presence of a separate attachment point of the igniter, inside which there is an afterburner made of heat-protective material. The increase in the diameter of the fastener and the structure caused by the coaxial location of the fastener over the afterburner does not allow to fully realize the potential advantages of combining the launch system and the pressure intake into a single unit.

The technical task of the present invention is to reduce the mass and dimensions of the design of the solid propellant rocket launcher system and the solid rocket motor pressure intake, solid rocket motor housing (cover), simplification of their manufacturing technology, increased reliability.

The essence of the invention "RDTT launch system" is that in the RDTT launch system containing pyrocartridges (squib) installed in the RTTT housing (cover), afterburner, igniter and its mount, the afterburner is threaded to form the ignitor mount. The igniter is mounted on the mount by means of a bottom having a fitting mating with the mount. A sleeve is installed outside the afterburner, a part of the internal channel of which is made with a diameter exceeding the external diameter of the fitting. The internal channel of the afterburner can be made with a variable diameter along the length, forming a receiver adjacent to the installation site of the squibs. The sleeve may be mated to the afterburner. In the bottom or fitting, holes can be made that communicate the internal channel of the afterburner with the internal cavity of the solid propellant rocket motor.

The essence of the invention, the “RTTT pressure intake” is that in the RTTT pressure intake, made around the RDTT start-up system and protecting its bushings, containing in the RTTT case (cover) of the telemetry measurement socket with channels connected to the internal cavity of the RTTT case, the screen covering the channels, the solid-propellant body (cover) and (or) the end face of the sleeve is provided with a step providing a gap between the solid-body body of the solid-propellant (cover) and the sleeve, the gap together with the screen forming an annular collector. The screen is made in the form of a hollow cylinder mounted coaxially to the sleeve, closing the gap, and channels go into the annular collector. A portion of the inner cylindrical surface of the screen can be made with an increased diameter and forms together with the sleeve an intake manifold, while in the sleeve from the intake manifold side blind radial holes are made, and longitudinal blind holes mating with them are made from the ring collector side. On the portion of the sleeve between the annular and intake manifolds, annular grooves intersecting the longitudinal holes can be made, and additional longitudinal holes are made between the longitudinal holes communicated with the radial holes. On the outer surface of the sleeve can be made ledge in contact with the reciprocal ledge made on the inner surface of the screen. The sleeve may be provided with a perforated wall covering the intake manifold.

The technical result in the launch system of solid propellant rocket engine is achieved by the implementation of the mount directly on the afterburner. The afterburner is made of metal, i.e. from the material possessing the strength necessary for the fastener. Despite the increased thermal conductivity of the metal, the exclusion (or minimization) of the heat-insulating plastic elements from the afterburner design is achieved by reducing the thermal effect from the combustion products in the internal channel of the afterburner due to the fact that, with a large elongation of the afterburner, its internal channel is a stagnant zone . The gas in the stagnant zone has a low mass and, accordingly, a small amount of heat. Part of this heat is transferred to the metal wall of the afterburner, providing some cooling of the stagnant zone. Moreover, the heating of the afterburner from the side of the stagnant zone is minimal due to the fact that the heat capacity of the metal afterburner is several orders of magnitude higher than the heat capacity of the portion of gas in the internal channel. Further, in the absence of circulation of combustion products, the stagnant zone performs a heat-insulating function, because it prevents the entry of new portions of hot gas into the internal channel of the afterburner. Those. significant convective heating of the exposed metal surfaces of the solid propellant rocket motor (cover) in the area of the pyro cartridge holders does not occur. At the same time, the end face of the afterburner is heated from the side of the fastening unit and heat transfer (heat distribution through the metal) to the side of the squibs. However, the length of the afterburner is sufficient so that in time its full warming would occur only at the end of the solid propellant rocket motor. The bottom, having a fitting mating with the attachment unit, can be either part of the solid propellant body (cover) or an igniter. The bottom provides a pair of different diameter elements - igniter and afterburner. The implementation of the receiver on the inner channel of the afterburner increases its free internal volume and somewhat enhances the effect of the stagnant zone. An increase in the free internal volume is advisable to reduce the pressure developed by the squibs when they are triggered, i.e. to reduce the load on the thread of both the squibs themselves and the igniter mount. The coupling of the sleeve with the afterburner on the thread ensures simplicity and reliability of the sleeve mounting, as well as the possibility of the sleeve fixing other elements, for example, the pressure intake screen. The implementation in the bottom or fitting of the holes communicating the internal channel of the afterburner 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 during the operation of the squibs.

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 RTTT start-up system due to the fact that for the annular manifold formed by the covered annular clearance screen, the radius of the annular collector does not exceed the radius of the sleeve (having, as shown , the smallest possible radius). Thus, the bosses in the solid propellant housing (cover) 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 start-up system and the pressure intake into a single unit is achieved by the fact that the solid-state solid-propellant housing (cover) 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 a minimum radius. In a single central boss, all the necessary sockets are made (for squibs and for STI sensors). The implementation of a separate intake manifold gas-coupled to the annular collector by means of blind radial holes made in the sleeve from the intake manifold side and longitudinal blind holes mating with them made from the annular collector side reduces the thermal effect on the STI sensors, i.e. improves the reliability of STI. Branches are formed in the sleeve, formed by additional longitudinal blind holes that are not connected with radial holes. Additional blind holes have a slightly smaller diameter compared to the main longitudinal blind holes and are combined with the main longitudinal blind holes by means of annular channels. Branches made in this way provide a more uniform distribution of pressure around the circumference of the annular manifold with sharp changes (or fluctuations) in the chamber pressure, i.e. the lack of circulation of combustion products and the performance of the STI when slagging part of the longitudinal blind holes, i.e. increase the reliability of STI. A step made on the outer surface of the sleeve in contact with the counter shoulder made on the inner surface of the screen provides a simple and reliable mechanical fixation of the screen when installing the sleeve on the afterburner by means of a thread. The perforated wall made on the sleeve and covering the intake manifold reduces the thermal effect on the STI sensors, i.e. improves the reliability of STI.

The proposed design of the RTTT pressure intake ensures minimizing the transverse dimensions of its main elements, and as a result, reducing their mass, simplifying manufacturing technology, improving the reliability of the structure.

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

The invention is illustrated by the following drawings:

figure 1 shows a view from the outside of the solid propellant rocket motor (cover) with the elements of the solid propellant rocket launcher system and the 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 the sleeve in isometry with a quarter cut.

The solid propellant rocket engine (RDTT) launch system comprises one or more pyro-cartridges 1 mounted in the boss 2 sockets made (see FIG. 2) in the solid propellant housing 3 (typically in the front cover of the housing 3). Sockets of the squib 1 are made on the side of the outer surface of the housing 3. On the side of the internal cavity of the body 3 of the solid propellant rocket (cover) a blind threaded hole 4 is made, which communicates with the sockets of the squibs 1. An afterburner tube 5 is installed in the blind threaded hole 4. The afterburner 5 has a through inner channel 6, made with a variable diameter along the length, forming from the side of the end mating with the solid-propellant body 3 (cover), receiver 7. From the side of the opposite end of the afterburner 5 on its outer surface is made in the form of a thread zel 8 mount. The igniter body 9 is made of quickly combustible material (foil or wicker basket). The igniter 9 is mounted on the mounting unit 8 by means of a bottom 10 having a nozzle 11 mating with the mounting unit 8. Holes 12 are made in the bottom 10 or the fitting 11, communicating the internal channel 6 of the afterburner 5 with the internal cavity of the solid-propellant body 1 of the solid-propellant rocket engine. Outside the afterburner 5, a sleeve 13 is installed, a part of the inner channel of which is made with a diameter greater than the outer diameter of the nozzle 11. The sleeve 13 is paired with the afterburner 5 by thread 14. The thread 14 can be made along the entire length of the wide part (above receiver 7) of the afterburner 5, i.e. it provides fixation of the afterburner 5 on the housing 1 of the solid propellant rocket motor (cover) and fixation of the sleeve 13 on the afterburner 5.

A solid fuel propellant rocket engine (RDTT) pressure intake is configured around the solid propellant launch system. It contains (see FIG. 3) in the boss 2 of the housing 1 of the solid-propellant rocket motor (cover) of the socket 15 of the telemetry measurement system (STI), gas-connected with the internal cavity of the housing 1 of solid rocket motor by means of channels 16. A screen 17 made in the form of a hollow is installed coaxially with the sleeve 13 the cylinder covering the annular manifold 18. The annular manifold 18 is formed by a gap 19 between the solid propellant housing 1 (cover) and the sleeve 13. In this case, the solid propellant housing 1 (cover), or the end face of the sleeve 13, is provided with a step 20, forming a gap 19. The step 20 can be performed on both of these elements (building ce 1 and the sleeve 13). Channels 16 go into the annular collector 18, communicating with the sockets 15 of the telemetric measurement system. Plugs 21 or STI sensors are installed in slots 15. A portion of the inner cylindrical surface of the screen 17 is made with an increased diameter and forms, together with the sleeve 13, an intake manifold 22. In the sleeve 13, blind radial holes 23 are made on the side of the intake manifold 22, and longitudinal blind holes associated with them are made on the side of the annular collector 18 (see Fig. 4 and Fig. 5). In the portion of the sleeve 13 between the annular 18 and the intake manifold 22, annular grooves 25 are made intersecting the longitudinal holes 26. Between the longitudinal holes 26 communicated with the radial holes 23, additional longitudinal holes 24 are made having a slightly smaller diameter compared to the main longitudinal holes 26 ( see figure 2 and figure 5). A step 27 is made on the outer surface of the sleeve 13, which is in contact with a mating step 28 made on the inner surface of the screen 17. The sleeve 13 is provided with a perforated wall 29 covering the intake manifold 22.

The device operates as follows. In the manufacture (assembly) and ground operation of the solid propellant rocket motor it is checked for leaks by pressurizing the internal cavity of the body 3 of the solid rocket motor with test pressure in one of two ways. According to the first option, the test pressure is supplied through the socket from the igniter 1, the internal channel 6 of the afterburner 5, holes 12. The pressure then passes through the intake manifold 22, radial 23 and longitudinal 26 holes to the sockets 15. Other sockets (socket for the second squib and socket 15) at the same time closed. Slots 15 are closed with either plugs 21 or STI sensors.

According to the second option, the test pressure is supplied through one of the sockets 15 through the longitudinal 26 and radial 23 openings into the internal cavity of the housing 3. Further, the pressure through the openings 12, the internal channel 6 of the afterburner 5 passes to the squibs 1. The rest of the sockets (second socket 15 and sockets for squibs) are closed.

Thus, when applying test pressure both through the socket from the squib 1, and through the socket 15, 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 15 of the STI.

If necessary, during the manufacture (assembly) and ground operation of solid propellant rocket motors, the internal cavity of the solid propellant chamber 3 is dried by repeatedly filling this cavity with dry gas. Filling is performed according to one of the two previously described options (i.e., both through the nest from the squib 1, and through the nest 15).

When conducting periodic tests of any tube 21 accidentally selected from a manufactured batch of solid propellant rocket motors (if information on the chamber pressure is not required by the aircraft control system, it is plugs 21 that are installed on the standard solid propellant rocket motor) are replaced by STI sensors. An engine equipped with either plugs 21 or STI sensors is ready for operation.

During the operation of the solid propellant rocket motor (i.e., until its launch), the loads applied to the igniter 9 are perceived by the bottom 10, the attachment unit 8, and a metal afterburner 5 firmly fixed to the solid propellant body 3 (cover). This ensures structural integrity.

When the command to start the engine, a current pulse is simultaneously supplied to both igniter 1 mounted on the solid propellant rocket motor. The initiation of the igniter 9 can occur when either one of the two squibs 1, or two at once, is triggered. The force of the flame from the squibs 1 through the afterburner 5 and the bottom 10 ignites the pyrotechnic composition of the igniter 9. Moreover, the pressure build-up in the through inner channel 6 of the afterburner 5 is moderate, because the presence of a receiver 7, which increases the internal volume, and holes 12, providing an additional consumption of combustion products from the squib 1, reduces the maximum pressure peak in the specified cavity. This provides a reduction in the load on the threaded connections on the squibs and the attachment unit 8 of the igniter 9. In the process of operation of the ignitor 9, the solid propellant rocket is started.

With further operation of the solid propellant rocket engine, the igniter body 9 burns out almost instantly, exposing the bottom 10 and the afterburner 5 located behind it. With a large elongation of the afterburner 5, its inner channel 6 and receiver 7 form a stagnant zone. The gas in the stagnant zone has a low mass and, accordingly, a small amount of heat. Part of this heat is transferred to the metal wall of the afterburner 5, providing some cooling of the stagnant zone. Moreover, the heating of the afterburner tube 5 from the side of the stagnant zone is minimal due to the fact that the heat capacity of the metal afterburner 5 is several orders of magnitude higher than the heat capacity of the portion of gas located in the internal channel 6 and receiver 7. Further, in the absence of circulation of the combustion products, the stagnant zone performs heat insulation function since it prevents the entry of new portions of hot gas into the inner channel 6 of the afterburner 5. That is significant convective heating of the open metal surfaces of the solid propellant housing 3 (cover) in the area of the nests of the squib 1 does not occur. At the same time, the end face of the afterburner tube 5 is heated from the side of the attachment unit 8 and heat transfer (heat propagation through the metal) to the side of the squibs 1. However, the length of the afterburner 5 is sufficient so that its full warming would only occur at the end of the solid propellant rocket motor operation.

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 15 STI sensors located in the sockets. Moreover, a system consisting of a perforated wall 29, an intake manifold 22, blind radial holes 23, longitudinal blind holes associated with them 26, an annular collector 18, channels 16, provides a significant reduction in thermal effects on STI sensors. Additional blind longitudinal holes 24, combined with the main longitudinal blind holes 26 through the annular channels 25, form branches that provide a more uniform distribution of pressure around the circumference of the annular manifold 18 with sudden changes (or fluctuations) in the chamber pressure. Thus, the absence of circulation of the combustion products and the operability of the STI when slagging part of the longitudinal blind holes 26 are ensured.

Feasibility study of the present invention, compared with the prototype, which selected the solid propellant rocket launcher system and the pressure intake made on the lid [Design of solid propellant rocket engines / Under total. ed. L.N. Lavrova. - M .: Mashinostroenie, 1993. - 215 p., Ill., Page 124, Fig. 2.55, diagram at the top left], consists in reducing the mass and dimensions of the design of the solid propellant rocket launcher system and the solid propellant pressure transducer, solid propellant rocket motor (RTTT) (simplification) technology for their manufacture, improving reliability.

Claims (9)

1. The rocket engine starting engine for solid fuel (RDTT) containing pyrocartridges (squib) installed in the body of the solid propellant rocket (cover), afterburner, igniter and its mount, characterized in that the afterburner is threaded to form the mount of the igniter, with this igniter is installed on the mount by means of a bottom having a fitting mating with the mount, and a sleeve is installed outside the afterburner, part of the inner channel of which is made with a diameter exceeding the outer diameter meter fitting. 2. The RDTT launch system according to claim 1, characterized in that the internal channel of the afterburner is made with a diameter that is variable in length, forming a receiver adjacent to the installation site of the squibs. 3. The RDTT launch system according to claim 1, characterized in that the sleeve is mated to the afterburner. 4. The RDTT launch system according to claim 1, characterized in that holes are made in the bottom or nozzle that communicate the internal channel of the afterburner with the internal cavity of the solid-state rocket engine. 5. The pressure intake of the solid propellant rocket engine (RDTT), made around the RDTT start-up system and the sleeves protecting it, containing the telemetry measurement sockets in the RDTT case (cover) with channels connected to the internal cavity of the RDTT case, a screen covering the channels, different the fact that the RDTT case (cover) and (or) the end face of the sleeve is equipped with a step providing a gap between the RDTT case (cover) and the sleeve, and the gap together with the screen forms an annular collector, while the screen is made in de installed coaxially with the hub of a hollow cylinder, closing the gap and out the channels in the annular manifold. 6. The RTTT pressure inlet according to claim 5, characterized in that the portion of the inner cylindrical surface of the screen is made with an increased diameter and forms an intake manifold together with the sleeve, and blind radial holes are made in the sleeve from the intake manifold side, and from the ring collector side longitudinal blind holes associated with them. 7. The RTTT pressure inlet according to claim 6, characterized in that in the portion of the sleeve between the annular and intake manifolds, annular grooves are made intersecting the longitudinal holes, and additional longitudinal holes are made between the longitudinal holes communicated with the radial holes. 8. The RTTT pressure inlet according to claim 5, characterized in that a step is made on the outer surface of the sleeve in contact with a reciprocal step made on the inner surface of the screen. 9. The RTTT pressure inlet according to claim 5, characterized in that the sleeve is provided with a perforated wall covering the intake manifold.

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