RU2412369C1 - Solid propellant rocket engine (versions) - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: solid propellant rocket engine comprises vessel, igniter, arranged in front or back part of engine, multi-nozzle unit concealed in vessel and insert charge of solid rocket propellant arranged in vessel and armoured along side surface and end. Multi-nozzle unit is arranged in front or back part of engine. Jacket made of elastic resilient material is installed tightly with interference onto charge at the side of armoured end, contacting with side armoured surface of charge, armoured end of change and engine bottom. Length of jacket makes 0.1…1.0 of length of armoured section of side surface of charge.

EFFECT: reduced smoke formation and increased reliability of solid propellant rocket engine with insert charge, due to elimination of combustion products overflow over armoured surface of charge.

2 cl, 4 dwg

Description

The invention relates to the field of rocket technology, and in particular to the design of a solid propellant rocket engine (solid propellant rocket engine) with an embedded armored solid rocket fuel (TRT) charge and can be used in the design, development and manufacture of solid propellant rocket engines.

Known designs of solid propellant rocket engines with embedded TPT charges armored over the outer surface according to patents: RU 2259495, RU 2164616, RU 2079689, RU 2282741, RU 2305201, RU 2233991, RU 2179989, US 3564845.

A common technical problem in the implementation of known constructions of this class is the difficulty of excluding during the operation of solid propellant rocket motors the flow of high-temperature gaseous products of combustion (PS) in the combustion chamber (CS) above the armored outer surface of the charge. As a rule, in such constructions of solid propellant rocket motors, in order to exclude PS overflows, a stagnant zone is organized (Figure 1), for example, by making an annular ledge on the armor plating (11) of the lateral surface of the charge (6) with the CS supporting surface between the ledge and the diaphragm (5) elastic gasket (8). The exclusion of PS flows in this case is ensured both by direct mechanical compression of the charge by the spring (4) to the gasket (8) and by the axial pressure difference in the CS (Δp = p ₁ -p ₂ ), where p ₁ is the pressure in the head of the engine, p ₂ - pressure in the pre-nozzle volume. This allows to exclude burnout of armor plating and abnormal (non-calculated) burning of the charge and the functioning of the solid propellant rocket engine.

A close analogue and quite effective of the known constructions of solid propellant solid propellant rocket motors of the type under consideration is the charge structure for the proprietary rocket engine according to US Pat. RU 2305201, IPC F02K 9/22, application RU 2005139878 dated 12/20/2005, providing (Figure 2) for firmly bonding an armor on the outer surface of the charge (6) TPT with a rear bottom (9) using (12) a liquid-viscosity polymerizable composition (ZhVPS) with the last monolithic filling of the lateral gap between the housing and the outer surface of the charge and with the exception of adhesion of the polymerizable composition to the outer surface of the charge. Such a design makes it possible, on the one hand, to ensure reliable fixation of the charge in the CS, on the other hand, to provide unloading of the nozzle end of the charge from the action of compressive stresses, and by practically filling the side gap of the air-to-air heater between the charge and the wall of the GG combustion chamber (RTTT), it is possible to eliminate TRT combustion products on the side (11) armor-plating charge. This, in turn, allows us to ensure the efficiency and reliability of both the charge and the solid propellant rocket motor. However, significant disadvantages of this design are the low manufacturability and the difficulty of its implementation when gluing the TRT charge into the housing, due to both significant labor, energy and time costs in the implementation of the polymerization of the liquid propellant polymer. As well as increased smoke generation during the operation of solid propellant rocket engines, due to the active thermal decomposition of both the armor plating of the charge and the composition filling the gap (GVPS).

Also known design of solid propellant rocket engine according to US Pat. RU 2079689 C1 (application 94004166 dated 02.08.1994, IPC F02K 9/08) according to the main similar features with the patented design and adopted by the authors as a prototype. The main disadvantage of the prototype is the lack of additional thermal protection of armor plating, which does not guarantee guaranteed to ensure the efficiency of the charge, including according to the required level of smoke formation.

An object of the invention is the development of the design of solid propellant rocket engines with an end or channel-end armored TRT charge on the side surface, which ensures the reliability of the solid propellant rocket engine due to the exclusion of PS overflows over the armored charge surface with a low level of smoke formation during operation of the rocket engine.

The technical result of the invention is to perform solid propellant rocket with a multi-nozzle block recessed into the engine housing and placed in the engine housing with an igniter and an auxiliary charge TRT channel-end combustion, armored on the side surface and rear end. In this case, the multi-nozzle block (10) of the engine and the igniter (2) of the charge are installed in the front of the engine (Figure 3), and from the side of the rear end of the charge, the cover (14) is made of an elastic material tightly, with an interference fit, in contact with the side armor coating ( 11) charge and its rear armored surface and rear end. The charge is made of ballistic or mixed solid rocket fuel, the armor plating (11) is made of thermoplastic armored material based on cellulose acetate or based on acrylate, polyurethane compositions, and the case (14) is made of resilient material, for example, rubber ИРП 3010. In this case, the igniter can be installed not in the front, but in the rear of the engine.

The technical result also consists in performing solid propellant rocket motors with a multi-nozzle unit and an igniter in the rear of the engine (Figure 4). At the same time, in the solid-propellant rocket motor housing there is a TRT plug-in end-face or channel-end combustion armored on the side surface and the front end, and from the front end side the charge is made of a cover made of elastic material tightly, with an interference fit, in contact with the side armored surface of the charge and its front armored end and front bottom of the engine. In this case, the charge is made of ballistic or mixed solid rocket fuel, the armor plating is made of thermoplastic armored material based on cellulose acetate or based on acrylate, polyurethane compositions, and the cover (14) is made of an elastic material, for example, rubber IPR 3010.

For patented structural variants of solid propellant rocket motors, the length of the cover along the side surface of the charge is within (0.1 ... 1.0) L _br , where L _bp is the length of the reserved area along the side surface of the charge.

The patented invention is illustrated in the figures.

Figure 1. The design of the solid propellant rocket chamber with a stagnant zone at the nozzle end of the charge

Figure 2. The design of the prototype (US Pat. RU 2305201).

Figure 3. Patented design of solid propellant rocket motors with a "recessed" multi-nozzle block and an igniter in the front of the engine.

Figure 4. Patented design of solid propellant rocket motors with a "recessed" multi-nozzle block and an igniter in the rear of the engine.

In figure 1, figure 2, figure 3, figure 4 are indicated:

1 - the front bottom of the engine

2 - igniter

3 - divider

4 - spring

5 - aperture

6 - charge TRT

7 - the body of the combustion chamber RDTT

8 - a sealing lining

9 - a back bottom of the engine

10 - nozzle block

11 - armor plating charge TRT

12 - ZhVPS

13 - nozzle cap

14 - elastically elastic cover.

The essence of the invention is to equip the charge armored on the side surface and end face (end or channel-end combustion) with an additional heat-protective coating in the form of a cover of elastic material, tightly, with an interference fit, mounted on the end and side surfaces of the charge. Due to the tight contact of the case with the armored charge, it is possible to provide additional thermal protection (armor protection) of the charge, to reduce the effect of high-temperature PS solid rocket fuel on the armor of the charge from the side of its outer surface. In the patented invention, the positive effect of the inventions according to US Pat. RU 2305201 (analogue) and US Pat. RU 2079689 (prototype). However, the implementation of the design of the analogue is accompanied by a complex technological process, including the preparation of a liquid-viscous polymerizable composition (ZhVPS), dosing it into the engine casing, immersing the charge in the casing by centering it, preliminary applying a release layer of the charge on the side armored surface and subsequent polymerisation of the bonding composition with increased temperature, and in the design of the prototype there is no external additional thermal protection of armor plating. In the patented design, the indicated disadvantages of the analogue and prototype are eliminated with a positive effect - increased reliability of the solid propellant rocket with a charge armored on the side surface with low smoke generation of the solid propellant rocket, which is very important when using the latter as part of guided missiles aimed at the target through the optical line of sight. The use of an elastic-elastic cover also allows, due to the selection of the thickness of its end section, to compensate for the free (assembly) axial clearance when equipped with a charge of the solid-state solid-propellant body, which allows limiting the movement of charge in the main body (solid-state solid rocket), reduces contact stresses on the support end of the charge and improves the operational reliability of the charge.

In the patented design of solid propellant rocket motors (Figure 3, Figure 4), when the igniter is activated (2) in a sealed compressor with nozzle plugs (13), which open at normalized pressure in the compressor (usually 30 ... 50 kgf / cm ² ), due to pressure waves from the igniter PS, the charge is tightly pressed through the end armored surface to the front (1) or rear bottom (9) of the solid propellant rocket through an elastic-elastic cover (14). This excludes PS flows after ignition of the TPT charge over the armored surfaces. The latter provides increased reliability of the solid propellant rocket engine and its low smoke generation. The indicated effect was experimentally confirmed during the fire bench tests of patented solid propellant rocket engine designs with measuring smoke generation parameters in accordance with US Pat. RU 2233991. The reduction in smoke generation is achieved by reducing the thermally degradable mass of the armor plating during the operation of the solid propellant rocket engine, due to the additional thermal protection of the armor plating of the side surface of the charge.

The implementation of the length of the cover is not less than 0.1L _br provides an effective seal on the end of the charge, which ensures the exclusion of gaseous PS flows over the side surface of the charge. The upper limit along the length of the cover provides reliable thermal protection of the side armor with low heat resistance (for example, based on cellulose acetate, patent RU 2179989). Intermediate values of L _bp can be used for armored coatings with a relatively high heat resistance, for example, based on smokeless additives (hydrazadicarbonamide and other acrylate compounds).

The patented technical solution is based on the following proposals:

1. TRT charges are made insert (not fastened to the solid propellant housing), with the possibility of free (but limited by the overall dimensions of the COP) movement inside the solid propellant rocket motor.

2. The TRT charges are small-sized (the tested range: length is up to 0.5 m, caliber is up to 0.15 m, weight is from 0.5 to 8.0 kg).

3. The charge igniter is carried out with a mass of sample (m _in ) sufficient to efficiently ignite the charge and create a pressure wave to move the charge to the bottom of the engine and tightly press it to it through a cover of elastic material (m _in = 1.0 ... 8.0 g smoky gunpowder).

4. The charges are made armored at the end and lateral surface.

5. A cover made of an elastic material is tightly mounted with an interference fit.

RTTT works as follows (Figure 3, Figure 4). After applying a pulse to the igniter, the igniter (2) is activated. Due to the pressure wave from the combustion products (PS) of the igniter, the charge (6) of the TPT moves to the front (1) or rear (9) bottom and is tightly pressed to it through an elastic cover (14). The igniter PS ignites the unarmored TPT charge surfaces. The resulting combustion products of the TRT charge, expiring through the nozzle block of the solid propellant rocket motor, provide thrust (R) of the engine. Moreover, in all the structural variants of the patented solid propellant solid propellant rocket, the armored end of the charge (with the cover (14) mounted on it (made of an elastic material)) mounted on the front (1) or rear (9) engine bottoms is provided, which ensures regular combustion of the TPT charge along equidistant surfaces with with the exception of burnout of the armored coating (11) and its reduced thermal decomposition (smoke formation) as a whole.

The positive effect of the invention is confirmed by the implementation of specific designs of solid propellant rocket motors, namely:

1. In the design of solid propellant rocket motors with a recessed nozzle block at the rear of the engine using a small-sized armored charge on the outer surface and front end of a slow- and medium-burning ballistic TRT with a burning speed of 2.5 to 12 mm / s at Рк = 40 kg / cm ² at a temperature (Т) of 20 ° С (Рк - pressure in the solid propellant solid propellant solid fuel).

2. In the design of the solid propellant rocket engine with a recessed nozzle block in the rear of the engine using a charge of quick-burning ballistic TPT, armored on the side surface and front end face weighing 5 ... 8 kg. TRT had a burning rate (U) at Pk = 100 kgf / cm ² , T = 20 ° C, U = 18 ... 24 mm / s.

3. In the design of the solid propellant rocket engine with a recessed nozzle block in front of the engine with a channel charge of slow-burning ballistic TPT, armored on the side surface and the rear end face with a mass of 1.5 ... 2.0 kg with a fuel burning rate at Pk = 40 kgf / cm ² , T = 20 ° C, U = 2.8-3.5 mm / s.

4. In the design of the solid propellant rocket motor with a recessed nozzle block solid propellant rocket engine in the front of the engine and an igniter located in the front of the engine with a channel charge, armored on the side surface, rear end and channel. In this case, the TRT charge had a burning rate of 18 ... 24 mm / s at Pk = 100 kgf / cm ² , T = 20 ° C.

Claims (2)

1. A rocket engine of solid fuel, comprising a housing, an igniter, a multi-nozzle block recessed into the housing, located at the front of the engine, and a plug-in solid rocket fuel of channel-end combustion, located in the housing and armored along the lateral surface and rear end, characterized in that the igniter is located in the front or rear of the engine, and the charge from the back of the armored end is fitted tightly with an interference fit of an elastic material in contact with the side armor the surface of the charge, the rear armored end of the charge and the rear bottom of the engine, while the length of the cover is (0.1 ... 1.0) L _br , where L _br is the length of the reserved portion of the side surface of the charge. 2. A solid propellant rocket engine comprising a housing, an ignitor located on the rear of the engine, a multi-nozzle block recessed into the housing and a solid rocket fuel charge placed in the housing, armored on the side surface and end face, characterized in that the solid rocket fuel charge is armored at the front end and represents the charge of the end or channel-end combustion, the multi-nozzle block is located on the rear of the engine, and the charge on the side of the front armored rtsa is fitted tightly with an interference fit of a case of elastic material in contact with the side armored surface of the charge, the front armored face of the charge and the front bottom of the engine, while the length of the cover is (0.1 ... 1.0) L _br , where L _bp is the length of the reserved area side surface charge.

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