RU2383764C1 - Solid propellant rocket engine - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: invention relates to rocketry and can be used in designing and producing solid propellant rocket engines, gas generators and solid propellant independent charges. Proposed engine comprises combustion chamber with nozzle unit, independent solid propellant charge and igniter. Nozzle unit is furnished with seal plug. Solid propellant charge represents a cylindrical monoblock with armored front face and side surface adjoining said front face. Igniter is arranged nearby rear unrestricted charge face. The side of aforesaid face features through channels arranged at an angle to charge lengthwise axis, channels outlets being located nearby extremity of the charge side armor.

EFFECT: reduced maximum pressure in starting operation, higher reliability, reduced weight.

5 dwg

Description

The invention relates to the field of rocket technology and can be used in the design, development and manufacture of solid propellant rocket engines (solid propellant rocket engines), gas generators (GG) and solid propellant solid propellant charges (TRT).

Known designs of solid propellant rocket engines (GT) with external charges TRT and charges TRT end combustion, including partially unarmoured side surface to the GG and marching stages of solid propellant rocket (US Pat. RU 2164616), or channel-end combustion (US Pat. RU 2259495) to accelerate -march steps of the solid propellant rocket engine.

Design (Figure 1) according to US Pat. RU 2164616 adopted by the authors for the prototype.

A feature of the prototype is the presence of a stagnant zone above the side armored charge surface.

A significant disadvantage of the prototype is the difficulty of providing the required parameters (P _max , P _min ) when the solid propellant rocket motor (GG) is in operation. Typically, in such systems, a high level of heat loss is realized when the stagnant zone (Figure 2) is filled with the igniter combustion products (PS) of the igniter, which causes significant difficulties when the solid propellant rocket generator (GG) reaches the operating mode at an acceptable maximum pressure level (P _max ). For the design of the prototype, a high probability of attenuation of the charge when entering the operating mode due to the presence of a significant "failure" in the diagram "pressure-time" (Figure 3).

An object of the invention is to provide the design of a solid propellant solid propellant rocket engine (GG) with an auxiliary charge TRT, which ensures reliable engine operation at a reduced maximum pressure (P _max ) in the solid propellant combustion chamber (GG) and a reduced “dip” in the pressure-time diagram.

The technical result of the invention consists in the performance of a solid propellant rocket engine containing a combustion chamber (KS) with a nozzle block equipped with a sealing plug, a fixed charge of solid rocket fuel placed in the combustion chamber in the form of a cylindrical monoblock, armored at the front end and partially along the side surface adjacent to the front end , and the igniter at the rear unarmored end. At the same time, through channels are made from the side of the rear unarmored end, at an angle to the longitudinal axis of the charge with the exit of the channel openings near the tip of the side armor coating (Figure 4).

The patented invention is illustrated by graphic materials.

Figure 1. Prototype design:

1 - combustion chamber

2 - nozzle block

3 - sealing plug

4 - charge TRT

5 - armor plating of the front end of the charge

6 - armor plating of the side surface of the charge

7 - rear unarmored end of charge

8 - igniter

9 - a sealing lining.

Figure 2. Charge ignition circuit for prototype construction.

Figure 3. Dependences "pressure-time" - p (τ) for the design of the prototype and patented design.

P - pressure in the COP, kgf / cm ³

P _max1 - the maximum pressure in the COP of the prototype, kgf / cm ²

P _max2 - maximum pressure in the COP of the patented design, kgf / cm ²

P _min1 - the minimum pressure in the COP of the prototype, kgf / cm ²

P _min2 - the minimum pressure in the COP of the patented design, kgf / cm ²

τ is the operating time, s.

Figure 4. Patented design:

1 - combustion chamber

2 - nozzle block

3 - sealing plug

4 - charge TRT

5 - armor plating of the front end of the charge

6 - armor plating of the side surface of the charge

7 - rear unarmored end of charge

8 - igniter

9 - a sealing lining

10 - through channels.

In figure 2, figure 5, the arrows show the movement of the PS in the initial period of ignition of the charge.

The essence of the invention is to reduce the influence of heat loss due to the presence of a stagnant zone in the solid propellant rocket igniter on the ignition of the charge as a whole and timely connection to the ignition and subsequent active gas intake from the side of the cylindrical unarmored part of the charge. Reducing the influence of heat loss on the nature of the pressure-time relationship according to the patented invention is achieved due to the active supply of PS igniter through the through channels directly into the stagnant zone, which ensures faster heating of the CS in the initial period and the subsequent decrease in the level of heat loss during the process of “connecting” to unarmored burning the cylindrical part of the charge TRT (Figure 5). This allows you to reduce the maximum pressure (P _max ) at the output of the solid propellant rocket engine to the operating mode and significantly reduce (Figure 3) the “failure” in the pressure-time diagram during the initial period of solid propellant rocket motor operation. In general, the proposed technical solution allows to improve, due to the optimization of ballistic characteristics, the weighted perfection of solid propellant rocket engines (GG), and therefore, to increase the efficiency of rocket systems.

Patented solid propellant solid propellant rocket motor includes (Figure 4) a combustion chamber (1), a nozzle block (2), a plug (3), a charge TPT (4) equipped with through channels (10), an ignitor (8) and a gasket (9).

RTTT works as follows (Figure 5). When a pulse is applied to the launch igniter, an igniter (8) is triggered, the PS of which is fed through the through channels (10) to the stagnant zone, simultaneously igniting the surfaces of the channels and the unarmored end (7) of the charge. The igniter PS and solid rocket fuel PS from flammable channels entering the “dead” part of the stagnant zone compensate for heat losses and, forming a gas flow from the “dead” part of the stagnant zone toward the nozzle block, ignite the lateral unarmored cylindrical surface of the charge (Fig. 5). With the specified ignition scheme of the patented solid-propellant solid-propellant structure, it is possible to obtain a pressure-time relationship p (τ) with a significant decrease in the initial pressure "dip" (Figure 3), which helps to increase the reliability of the solid-propellant rocket motor and achieve more optimal ballistic characteristics.

An example of the practical implementation of the patented design of solid propellant rocket motors (Figure 4).

The solid propellant solid propellant solid propellant solid propellant rocket contains a face combustion TRT charge made of slow-burning fuel with a burning rate at T = 20 ° C, P = 100 kgf / cm ² - 2 mm / s. Charge Dimensions:

- length - 130 mm,

- outer (fuel) diameter - 100 mm,

- the length of the unarmored cylindrical portion (L) of the charge is 30 mm.

The charge is armored at the front end and partially along the side surface. From the side of the unarmored end, four through channels with a diameter of 6 mm each were made with the exit of the channel openings to the tip of the armored coating (at the entrance to the "stagnant zone"). As an igniter, we used a sample of smoke gunpowder DRP-2 weighing 7 tons, placed at the rear unarmored end of the charge. The sealing plug provided pressure up to opening - 20 ... 30 kgf / cm ² . The dependence P (τ) for this example is shown in Fig.3.

Claims (1)

A solid propellant rocket engine containing a combustion chamber with a nozzle block equipped with a sealing plug, a solid rocket propellant charge placed in the combustion chamber in the form of a cylindrical monoblock armored at the front end and a side surface adjacent to the front end, and an igniter at the rear unarmored end of the charge , characterized in that from the side of the unarmored end of the charge through channels are made at an angle to the longitudinal axis of the charge, the outlet openings of which are located in near the tip of the side armor plating charge.

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