RU2377431C2 - Solid propellant ascent engine - Google Patents

Abstract

FIELD: aerospace engineering, engines and pumps.

SUBSTANCE: invention relates to ascent solid propellant engines wherein power conversion period is limited by second hundredth and thousandth. Proposed engine comprises shell and engine gasdynamic section comprising central perforated tube with bundle of solid propellant disk-type elements arranged thereabout and in perforated combustion chamber, igniter and outer flow annular nozzle. Combustion chamber features cylindrical casing with inclined tangential slots with flow section area decreasing along combustion chamber length from section to section. Solid propellant disk-type elements feature equal size and integrated into a charge with the help of powder supports glued onto their face surfaces. Powder supports feature burning hill thickness equal to half the thickness of disk-type element burning hill.

EFFECT: more complete combustion of powder charges, increased specific thrust pulse.

2 dwg

Description

The invention relates to the field of solid rocket starting rocket engines (SRDTT), in which the chemical energy of the powder charge is converted into the thermal energy of gases, and then into the kinetic energy of the expiring gas jet, in particular to SRDTT, in which the energy conversion time is determined in hundredths and thousandths in fractions of a second.

A known charge of solid rocket fuel containing a lid filled with an adhesive composition in which pins with channel powder checkers are mounted, each pin is glued into the channel powder checker and made with an adhesive ring lock on a cylindrical head with a diameter equal to the diameter of the checker channel, and a spring leg with a support rigidly installed in the lid with an air gap between the channel powder checkers and the adhesive (see patent of the Russian Federation RU 2255240 C1 of June 27, 2005, IPC F02K 9/32). A significant disadvantage of solid rocket propellant charges consisting of cylindrical elements with a central hole is the low density of the combustion chamber. Other disadvantages of this design of powder charges include the possibility of its destruction due to the pressure drop between the checker channel and the gaps formed by the outer surfaces of the powder checkers, and the presence of a significant erosion effect, which at the final stage of the engine’s operation leads to burnout of the powder checkers due to their divergence , their destruction and removal of a certain part of the powder particles from the engine chamber. These circumstances, due to operational requirements for gunpowder, i.e. the minimum permissible thickness of the burning vault of the powder block, increase the overall dimensions of the engine and determine the minimum possible operating time of the solid propellant rocket engine, and the release of the destroyed powder particles causes significant losses in the specific impulse of the traction force and the spread of the engine operating time. The same circumstances practically exclude the use of SRDTT in constructions, except for pyroxylin and other powders (ballistic, mixed) due to their low specific strength to destructive loads.

These disadvantages are largely eliminated in the design of a solid rocket starting rocket engine containing a shell (transport and launch container) and a gas-dynamic path, including a charge consisting of a bundle of solid-fuel disk elements, an igniter and an annular external flow nozzle, while the gas-dynamic engine path also includes a central perforated tube, around which there is a bundle of solid fuel disk elements placed in a perforated chamber (see GB 158 6109 A dated 03/18/1981, IPC F02K 9/10). However, this design of the charge of solid rocket fuel, eliminating the erosive combustion of gunpowder, also has several disadvantages. The manufacture of a cone-shaped housing of the combustion chamber implies the use of disk elements of solid fuel of different diameters, which in turn leads to a different initial combustion surface of the disk elements. In view of burnout of the disk elements due to their divergence, the amount of the destroyed fuel particles will depend on the size of the initial combustion surface of the disk element, and at the final stage of the functioning of the solid-propellant rocket engine, the position of the particles at the time of their formation is important. Implemented radial-axial outflow of combustion products in the proposed design SRTT prototype suggests the need to provide the required residence time of the fuel particles within the engine path, limited in length to the cross section of the confuser of the annular nozzle. Taking into account the design of the cone-shaped housing of the combustion chamber and the execution of radial holes of the same area on it, and also taking into account the inertial forces arising from the operation of the proposed design of the solid propellant rocket engine, it becomes obvious that the destroyed fuel particles after “forcing” through the radial holes of the chamber can leave it completely not burned out. A certain amount of solid fuel particles removed from the SRDT chamber causes certain losses of the total impulse of the thrust force and the spread of the ballistic parameters, such as the internal chamber pressure and the total operating time of the engine, which in turn will lead to a spread in the main ballistic characteristic - the full impulse of the thrust force.

The objective of the invention is to increase the completeness of combustion of the charge of solid rocket fuel and reduce the variation in the operating time of the solid propellant rocket engine due to the improvement of the intracameral workflow by organizing the tangential-radial outflow of combustion products along the gas-dynamic path of the engine.

This task is achieved by the fact that the proposed structural changes of the well-known solid rocket starting rocket engine containing a shell and a gas dynamic path of the engine, which includes a central perforated tube, around which there is a bunch of solid fuel disk elements placed in a perforated combustion chamber, an igniter and an annular external nozzle flow around.

Distinctive features is that the combustion chamber has a cylindrical body in which inclined tangential slots are made, while the area of their passage sections decreases along the length of the combustion chamber from section to section, the disk elements of solid fuel are made of the same size and collected into a charge using glued on their end surfaces of powder supports having a thickness of the burning arch equal to half the thickness of the burning arch of the disk element.

The essence of the invention lies in the fact that the claimed design of the SRDTT provides a more complete combustion of the charge of the disk powder elements within the gas-dynamic path by increasing the residence time of the burning solid fuel particles formed after the destruction of the charge, achieved by creating a radially tangential component of the flow of the effluent combustion products. The selection of the geometric characteristics of the proposed design of solid propellant solid propellant rocket engines and solid fuel disk elements allows, using the swirling flow of combustion products, to increase the completeness of the energy potential of the charge of solid fuel.

Brief Description of the Drawings

Figure 1 shows the proposed design of the SRDTT, and figure 2 is a diagram of the installation of intermediate supports on the disk element of solid fuel, where

1 - shell;

2 - front end wall;

3 - housing perforated combustion chamber;

4 - rear end wall;

5 - inclined tangential slits;

6 - annular nozzle of the external flow around;

7 - gas-dynamic channel;

8 - main afterburner membrane;

9 - disk element of solid fuel;

10 - radial distances;

11 - Central perforated tube;

12 - radial holes;

13 - powder support;

14 - igniter composition;

15 - electric igniter;

16 - igniter afterburner membrane.

The proposed design, shown in Fig. 1, is a drawing of a solid propellant combustion chamber, the rocket part of which is shown in section along with a shell 1. The solid propellant rocket engine consists of a perforated combustion chamber formed by a cylindrical body of a perforated combustion chamber 3, as well as a front 2 and rear end wall 4. The front end wall 2 is located in the shell 1 and is the master device. The housing of the perforated combustion chamber 3 has a plurality of inclined tangential slots 5, while the area of their passage sections decreases along the length of the housing from the front to the rear end wall. The case of a perforated combustion chamber is the main structural element that ensures the stability and strength of the structure as a whole. An annular external flow nozzle 6 is attached to the rear end wall. The gas-dynamic channel 7 is formed by the outer surfaces of the housing of the perforated combustion chamber 3, the annular external flow around the nozzle 6 and the inner surfaces of the front end wall 2 and the shell 1. A bundle of hard disk elements is located in the internal volume of the perforated combustion chamber fuel 9 with radial distances 10 between them, providing a radial flow of combustion products to inclined tangential slots 5. Disk element Made of identical dimensions, are mounted on a central perforated tube 11 with the radial distances 10, secured by means of stickers to their end faces 13 poles powder having a burning arch thickness equal to half the thickness of the burning arch disk member, which installation diagram is represented in Figure 2. The Central perforated tube 11 serves as an additional support between the end walls 2 and 4 and also provides structural strength, in addition, it is a chamber for the igniter composition 14, which is a sample of smoky gun powder. An electric igniter 15 is arranged in the inner cavity of the central perforated tube, which is designed to transmit a fire pulse to the igniter composition 14. The perforated tube has radial holes 12, which ensure forced flow of the combustion products of the igniter composition and ignition of solid fuel disk elements. To ensure the necessary gas pressure from the combustion of the igniter composition, an afterburner membrane 16 was used, with the same purpose the main afterburner membrane 8 was used to ensure stable ignition of the disk elements. Aluminum foil is used to produce both the ignitor and the main afterburner membranes.

The functioning of the proposed SRDTT is as follows: when the electric igniter 15 is activated, the igniter composition 14 is ignited. When the forcing pressure is reached, the igniter afterburner membrane 16 burns out and the igniter combustion products enter the perforated combustion chamber through the radial holes 12 through the internal cavities of the radial distances 10, as a result of which the disk elements of solid fuel 9 are ignited along their end surfaces. When the boost pressure is reached in the volume of the perforated combustion chamber, the main afterburner membrane 8 burns out and the solid fuel combustion products, spinning, expire through the inclined tangential slots 5 into the gas-dynamic channel 7.

As shown by the authors of experimental studies, by varying the angles of inclination of the tangential slots, the swirl directions and the ratio of the areas of the passage sections of the inclined tangential slots, it is possible to increase the specific impulse of the traction force to 15% with equal masses of charges and to reduce the spread in the operating time of the solid-state solid-propellant rocket engine by 1.5 times comparison with the prototype, which confirms the possibility of ensuring high efficiency of the proposed design SRDTT.

Claims (1)

A solid fuel starting rocket engine containing a shell and a gas dynamic path of the engine, which includes a central perforated tube, around which there is a bundle of solid fuel disk elements placed in a perforated combustion chamber, an ignitor and an annular external flow nozzle, characterized in that the combustion chamber has a cylindrical a housing in which inclined tangential slots are made, while the area of their bore sections decreases along the length of the combustion chamber from section to section , Disc elements made of solid fuel and of the same size are collected in the charge applied on via their end surfaces of supports powder having a burning arch thickness equal to half the thickness of the burning arch disc member.

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