RU2704058C9 - Separating bottom of multi-impulse solid propellant rocket engine - Google Patents

Abstract

FIELD: engines and pumps.SUBSTANCE: separating bottom of multi-impulse solid-propellant rocket engine comprises housing, launching and sustainer chambers with powder channel charges. Bottom is monolithic with groups of perforated holes symmetrically located relative to the longitudinal axis and closed by flat thin-walled membranes tightly fixed on the dividing bottom and having heat-resistant coating. Groups of perforated holes in monolithic separating bottom are arranged arbitrarily relative to channels (slots) of sustainer chamber, and size of holes is selected from ratios protected by present invention.EFFECT: invention allows improving strength and reliability of rocket engine and reducing labor intensity of manufacturing.1 cl, 5 dwg

Description

The invention relates to the field of rocket technology, namely to multi-pulse solid propellant rocket engines, and can be used to create rockets.

Separation bottom as part of a multi-pulse solid fuel rocket engine, in particular a two-pulse one, consisting of a cylindrical body, launch and sustainer chambers with powder channel charges and nozzles. The separation bottom is located between the chambers and is a monolithic design with groups of holes symmetrically located relative to the longitudinal axis and closed with flat thin-walled membranes, hermetically fixed to the separation bottom and having a heat-protective coating. The groups of holes are located randomly relative to the charge channels of the marching stage, and their size is selected from the ratios indicated in the present invention, confirmed by bench tests.

The closest in technical essence and the achieved positive effect is a dual-mode solid fuel rocket engine (Russian patent No. 2390646 IPC F02K 9/28 (2006/01), containing a cylindrical body, launch and marching chambers with powder channel charges, a separation bottom, a bottom opening device and the nozzle. The opening device is made in the form of spherical perforated plugs symmetrically located relative to the longitudinal axis of the engine, curved towards the marching chamber, and from the start side oh chamber stub interacts with the thin-walled diaphragm repeats the shape of the plug sealingly clamped peripherally on the plug, and having a thermal barrier coating, wherein the plugs are disposed opposite channels sustainer propellant charge chamber.

The disadvantage of this device is the following.

To obtain optimal ballistic characteristics at the start, usually use the maximum possible operating pressure in the combustion chamber of the start mode. At this pressure, the stresses arising in the design of the separation bottom are close to permissible, taking into account the safety margin.

A number of main holes are made in the dividing bottom, into which then perforated plugs are inserted, which provide sealing of the launch chamber. In the case of using the charge of the marching chamber, which has a complex shape, for example, with the number of channels (slots) more than 6, a corresponding increase in the number of holes for plugs located opposite the channels (slots) of the charge leads to a decrease in the strength of the separation bottom in the places of the jumpers between the holes due to their thinning and deformation of the bottom towards the marching chamber by an unacceptable value, which significantly reduces the reliability of the engine.

The aim of the invention is to increase the strength and reliability of the engine.

This goal is achieved by the fact that the separation bottom as part of a multi-pulse solid fuel rocket engine containing a cylindrical body, a launch and marching chamber with powder channel charges and a nozzle is made monolithic with groups of holes no more than six, located symmetrically relative to the longitudinal axis of the engine, closed flat thin-walled membranes, hermetically fixed around the periphery of the holes and having a heat-shielding coating, while groups of holes can be located in the openings arbitrary position relative to the channels (slots) of the sustainer charge, and their total size at the beginning of the operation of the sustainer chamber and the thickness of the membrane are selected from the ratios:

n⋅S _holes ≥1,45S _kr

2δ _min ⋅τ _cf. ≥R _ks ⋅R⋅K,

where n is the total number of holes;

S _holes - hole area;

S _cr - the area of the critical section of the nozzle;

δ _{min - the} minimum thickness of the membrane;

τ _cf - ultimate stress on the slice of the membrane material;

R _ks - pressure in the launch chamber;

R is the radius of the hole;

K is the safety factor.

The invention is illustrated by drawings, where in FIG. 1 shows a general view of the device, FIG. 2 - possible arrangement of channels (slots) in the charge of the sustainer chamber, in FIG. 3 - the location of the holes in the bottom, in FIG. 4 - bottom structure, FIG. 5 - membrane with notches.

The device (Fig. 1) contains a combustion chamber 1 with a solid fuel charge 2 providing a starting mode, a combustion chamber 3 with a solid fuel charge 4 providing a marching mode. In the charge 4, slotted channels 6 are made (Fig. 2).

The start (1) and mid-flight (3) chambers are separated by the bottom 5. The dividing metal bottom 5 is protected by thermal insulation 12 from the side of the launch chamber and thermal insulation 13 from the side of the mid-chamber (Fig. 4). In the bottom (Fig. 3) there are six groups of holes, the total bore of which is 1.45 S _cr , which ensures subsonic outflow of combustion products and reliable operation of the marching stage. From the side of the launch chamber, the groups of holes 9 are closed by thin-walled flat membranes 7 with thermal insulation 8 (Fig. 4).

A radial lobe notch 11 is applied to the membrane 7 (Fig. 5), which ensures its stable opening when the specified pressure in the marching chamber is reached.

For the expiration of combustion products from the launch and mid-flight chambers, one nozzle 10 with a critical cross-sectional area S _{cr is used} .

The proposed device operates as follows. The start and marching cameras are switched on sequentially. When the start chamber is turned on, the pressure of the combustion products tends to push a thin-walled membrane through the holes in the separation bottom, creating maximum stresses around the perimeter of the hole.

When the launch chamber is used to prevent the opening of the calculated membrane, the load from the pressure on the local portion of the membrane corresponding to the area of one hole is accepted, which allows the use of a thin-walled flat membrane for sealing.

When the sustainer chamber is switched on, the combustion products, penetrating through the holes, act on the membrane over its entire area, open it and, penetrating the launch chamber, expire through the nozzle.

The use of radial notches in the membrane allows for their stable opening when a certain pressure is reached.

By bench tests, it was confirmed that the total passage area in the separation bottom, which is 1.45 S _cr , ensures subsonic outflow of combustion products during the operation of the sustainer chamber and reliable operation of the engine.

The proposed technical solution allows to increase the reliability and strength of the engine by applying simple design solutions.

Claims (11)

Separation bottom of a multipulse solid propellant rocket engine containing a cylindrical body, launch and sustainer chambers with powder channel charges, a nozzle characterized in that the separation bottom is made monolithic with groups of holes, no more than six groups, located symmetrically relative to the longitudinal axis of the engine, closed with flat thin-walled membranes hermetically fixed around the periphery of the holes and having a heat-shielding coating, while the groups of holes are arranged in random th position relative to the channels (slots) of the sustainer charge, and their total size at the beginning of the operation of the sustainer chamber and the membrane thickness are selected from the relations: nS _holes ≥1,45S _kr 2δ _min ⋅τ _cf. ≥R _ks ⋅R⋅K, where n is the total number of holes; S _holes - hole area; S _cr - the area of the critical section of the nozzle; δ _min ^{- the} minimum thickness of the membrane; τ _cf - ultimate stress on the slice of the membrane material; R _ks - pressure in the launch chamber; R is the radius of the hole; K is the safety factor.

Images