RU2620613C1 - Rocket engine of rocket-assisted projectile - Google Patents

Abstract

FIELD: engines and pumps.

SUBSTANCE: rocket engine of rocket-assisted projectile comprises a combustion chamber with a charge of solid fuel nozzle, and the nozzle cap initiator. In the critical section of the nozzle is set breakthrough membrane. The plug includes a base and a cover fixed on the basis of a hollow cylindrical cup with perforated bottom by a membrane installed at the throat of the nozzle. At the base of the plug and the bottom of the glass made coaxial holes in which the rod is installed with the possibility of longitudinal movement. The stock has a sharpened tip from the membrane, a conical thickening of the part of the base cap, combined with the conical recess in the base and cut the flange sandwiched between the base and the plug cover. On the inside the glass rod mounted console, and between the glass and the bottom bracket mounted coil spring, covering the stem. Pyrotechnic initiator primary linkage includes an igniter placed between the cup bottom and the membrane, and at least two igniters kaplyuley-established on the basis of plugs and associated with strikers fixed to the console. Cover nozzle stub located at the outlet section of the nozzle and fastened by means of rolling with the outer side and the central portion of the lid is a hole with a diameter equal to the diameter of the tapered stem thickening. The magnitude of free volume combustion chamber defined by an algebraic expression to be protected by the present invention.

EFFECT: invention allows for reliable autonomous ignition charge of solid fuel, which is independent from the influence of powder gases and propellant pressure relief at the departure of the nozzle cap.

4 dwg, 1 tbl

Description

The invention relates to the field of rocket technology, in particular to rocket engines of active rockets (ARS), launched from the barrel of an artillery gun, and can be used in the development of rocket engines included in the flight path of the APC.

A feature of the internal ballistic of the ARS is that the pressure of the propellant propellant gases in the gun barrel is several thousand atmospheres (artillery pressure range), and the working pressure in the combustion chamber of a rocket engine is several tens of atmospheres (rocket pressure range) [1]. Exposure to high-pressure gases can lead to deformation of the surface of the charge of a rocket engine, disruption of internal ballistics and destruction of the engine. To protect the propellant’s fuel charge from ignition in the gun’s barrel, various types of plugs installed in the nozzle block of the rocket engine are used in the APC designs.

Known rocket engine artillery shell [2], containing a housing with a charge of solid fuel and a nozzle, a plug installed in it with an annular protrusion in contact with the outer surface of the nozzle, fixing its element, an ignitor and a knife-type stabilizer block. The igniter is located on the plug in the conical body and is equipped with a moderator mounted on the nozzle exit side. The stabilizers overlap the end of the moderator, and the locking element is made in the form of a tube, one end of which is fastened to the nozzle in its critical section, and the other end is provided with protrusions covering the igniter body.

Known rocket engine of solid fuel of an active-reactive projectile [3], comprising a housing with a nozzle block and a charge of solid fuel, a plug installed in the nozzle block with a central channel in which the pyro-retarder is placed, and channels in its bottom facing the nozzle exit, located in housing from the side of the pyro-moderator igniter. Moreover, the channels in the bottom of the plug are made radial, at the entrance of which curvilinear recesses are formed.

Known rocket engine artillery shell [4], containing a combustion chamber with a charge of solid fuel, a nozzle, a pyrotechnic initiator delayed action installed in the nozzle plug channel. The engine additionally contains a perforated disk and a ball, and in the plug from the nozzle exit side the channel is made with expansion towards the initiator. The ball is placed in the cavity of the channel, and a perforated disk is installed between the channel and the initiator, and the holes in the disk are made with a diameter smaller than the diameter of the ball.

Known active-rocket projectile [5], which contains a rocket engine of solid fuel, a nozzle plug with a pyro-moderator, a charge of solid fuel and an ignitor. A semi-closed cavity is made at the rear end of the projectile, while the moderator is buried in this cavity. At the end of the moderator from the rear end of the projectile mounted throttle washer made with at least one transverse diametrical groove. The width of the groove is made smaller than the diameter of the throttle washer.

Closest to the technical solution to the claimed invention is a rocket engine of an artillery shell [6]. The engine contains a housing with a solid fuel charge and a nozzle, blocked in the output part by a conical plug with an igniter and moderator. At the end of the wall of the plug facing the charge, an annular protrusion is made, on which a membrane in the form of a cup with a flanging is installed with a gap relative to the end filled with a sealing compound.

A disadvantage of the known technical solutions is, in addition to the complexity of the design of most of them, the presence of an additional solid fuel charge (moderator). The retarder is initiated by propellant propellant gases in the gun’s barrel and, in turn, initiates the solid propellant charge of the mid-flight rocket engine after the APC takes off from the barrel. The presence of a moderator reduces the reliability of launching a rocket engine due to the possibility of developing an unsteady mode of combustion of the moderator (up to its extinction) under conditions of a sharp pressure relief when the projectile leaves the barrel.

The technical result of the present invention is to increase the reliability of the initiation and operation of the mid-flight rocket engine of an active-rocket projectile.

The technical result of the invention is achieved by the fact that a rocket engine of an active-rocket projectile is developed comprising a combustion chamber with a charge of solid fuel, a nozzle, a pyrotechnic initiator and a nozzle plug. A breakthrough membrane is installed in the critical section of the nozzle; the plug consists of a base, a cover, and a hollow cylindrical glass fixed to the base with a perforated bottom on the side of the membrane. At the base of the plug and the bottom of the glass, coaxial holes are made in which the stem is mounted with the possibility of its longitudinal movement. The stem has a pointed tip on the membrane side, a conical thickening on the side of the base of the plug associated with a conical recess in the base, and a cut flange sandwiched between the base and the cover of the plug. A cantilever is fixed on the stem inside the cup; a cylindrical spring is installed in the cup between its bottom and the console, covering the stock. The pyrotechnic initiator consists of a sample of the main igniter located between the bottom of the glass and the membrane, and at least two igniter caps mounted on the base of the plug and paired with drums mounted on the console. The nozzle cap cover is located in the outlet section of the nozzle and is fixed by rolling from its outer side, a hole is made in the central part of the cap whose diameter is equal to the diameter of the conical thickening of the rod. In the combustion chamber there is a free volume between the charge of solid fuel and the inlet section of the nozzle, and the free volume of the combustion chamber corresponds to the inequality

,

,

where V is the free volume of the combustion chamber, m ³ ;

;

;

k is the adiabatic exponent of the solid fuel combustion products;

- приведенная сила топлива, м²/с²;

- reduced fuel power, m ² / s ² ;

R is the gas constant of the solid fuel combustion products, J / (kg⋅K);

T _p - adiabatic combustion temperature of solid fuel at constant pressure, K;

ρ _m - density of the solid fuel, kg / m ^3;

S _t - surface area of the combustion of a charge of solid fuel, m ² ;

u ₁ - burning rate of solid fuel at atmospheric pressure, m / s;

p _{to the} opening pressure of the nozzle plug, Pa;

p ₁ = 0.1 MPa - atmospheric pressure;

n = p _to / p ₀ - pressure ratio;

p ₀ - pressure in the combustion chamber at the marching mode of operation of the rocket engine, Pa;

ν is the exponent in the law of the burning rate of solid fuel;

The invention is illustrated by the scheme of the rocket engine of an active-rocket projectile (Fig. 1). The engine comprises a housing 1 with a charge of solid fuel 2 and a nozzle blocked in a critical section 4 by a breakthrough membrane 22. In the diffuser of the nozzle 6, the base of the plug 8 is installed, which is fastened to the outlet part of the diffuser by a cover 13 with a hole 12 by rolling 7. On the base of the plug 8, a cup 17 is fixed on the charge side 2, along the axis of which a movable rod 19 with a console 16 mounted on it is mounted. On the side of the base of the plug 8, drums 15 are mounted on the console 16, and on the charge side 2, the console, through the spring 18, rests on the bottom of the cup 17. On the base of the plug 8, igniter caps 14 are placed opposite the impactors 15. Between the bottom of the cup 17 and the breakthrough membrane 22, a sample of the main igniter 20 is placed. Perforations 5 in the bottom of the cup 17 connect the inner cavity of the cup with the igniter 20. A pointed tip 21 is made on the rod 19 from the side of the breakthrough membrane 22, the conical base 11 from the base side of the plug 8, conjugated with a conical cavity 9 in the base of the plug 8, and a shear flange 10, which is sandwiched between the base of the plug 8 and the cover 13.

The rocket engine (solid propellant rocket engine) active-rocket projectile operates as follows. When the projectile moves in the gun’s barrel under the influence of high pressure of the propellant propellant’s gas, the protruding edge of the flange 10 is cut off and the stem 19 moves towards the membrane 22. In this case, the conical base 11 of the rod 19 mates with the conical recess 9 at the base of the plug 8, preventing the breakthrough of the powder gases into the internal cavity of the cup 17. With the movement of the rod 19, the spring 18 is compressed, and the tip 21 breaks through the membrane 22.

. При р_к≤р₀ сопловая заглушка вскрывается раньше момента воспламенения заряда, вследствие чего продукты сгорания воспламенителя сбрасываются через сопло.After the projectile leaves the gun at time t = 0 (Fig. 2), the pressure of the powder gases at the bottom of the projectile sharply decreases, while the rod 19 moves backward under the action of the spring 18, opening a hole in the breakthrough membrane 22 and acting on the igniter caps with impactors 15 14. Force of the flame from the primer 14 through the perforations 5 in the glass core 17 ignites the igniter 20. The igniter 20 of the combustion products through the opening 22 in the membrane disruptive fed into the combustion chamber of the rocket engine 3 and at time t _to a value of the pressure in the combustion chamber p _in (Fig. 2) ignites the charge of solid fuel 2. When the specified pressure p _k is reached in the combustion chamber at time t _k (Fig. 2), the rolling 7 of the cover 13 is cut off and the plug is ejected from the nozzle block. In the combustion chamber at time t _k (Fig. 2), a pressure is released from p _k to the operating pressure p _{0 of the} cruise mode of the engine, which is set at time t ₀ . From FIG. 2 it follows that to ensure reliable ignition of the solid fuel charge, the opening pressure of the nozzle plug p _to should be significantly higher than the working pressure p ₀ sustainer mode of the engine:

. At p _k ≤ p _{0, the} nozzle plug opens before the moment of ignition of the charge, as a result of which the products of combustion of the igniter are discharged through the nozzle.

The achievement of the positive effect of the invention is provided by the following factors.

1. The use of a shock mechanism with an igniter capsule to initiate ignition of the igniter instead of a moderator transmitting combustion from powder gases increases the reliability of the device, because isolates the igniter composition from high-pressure gases in the gun barrel and eliminates the possibility of failure due to the extinction of the moderator during sudden pressure fluctuations and during overloads.

2. Placing the rod with a shock mechanism and a return spring in the glass ensures their free movement and transfer of flame from the igniter capsule to the main igniter composition.

3. The use of a rod with a pointed tip makes it easy to destroy the membrane during its translational motion and provides free flow of igniter gases through the rupture of the membrane when the rod is displaced in the opposite direction under the action of a spring.

4. The conical thickening on the stem from the side of the base of the plug, coupled with a conical recess in the base of the plug, provides a tight fit of the stem in the base of the plug during its forward movement and prevents breakthrough of powder gases into the igniter body.

5. The installation of a breakthrough membrane in the critical section of the nozzle and its preliminary opening by the rod tip eliminates the formation of fragments of the membrane, allows you to simply and reliably seal the main igniter composition.

6. The execution of the flange at the end of the rod, facing the base of the plug and sandwiched between the base of the plug and the cover, allows you to rigidly fix the stem from accidental movements and seal the inner cavity of the plug.

7. To determine the free volume of the solid propellant combustion chamber, we consider the equation of energy conservation in the solid propellant chamber during pressure relief [7]:

where p _to - the pressure in the combustion chamber corresponding to the opening pressure of the nozzle plugs (cutting rolling of the lid);

t is the time;

k is the adiabatic exponent of the TRT combustion products;

- приведенная сила топлива;

- reduced fuel power;

R is the gas constant of the combustion products of the TRT;

T _p - TPT adiabatic combustion temperature at a constant pressure;

V is the free volume of the combustion chamber;

G ₊ - mass second gas inlet during combustion of TPT;

G _- - mass second consumption of TPT combustion products through the nozzle.

The condition for quenching the solid fuel charge when the pressure is released at the time of opening the nozzle plug is determined by the inequality [8]

where parameter B depends on the type of TRT:

At = 10 s ^-1 - for ballistic TRT;

B = 120 s ^-1 - for mixed TRT.

The values of the gas inlet G ₊ and the flow rate G _- combustion products are determined by the equations [7]:

where ρ _t - density TRT;

S _t - surface area of the combustion of a charge of solid fuel;

u ₁ - the rate of combustion of TRT at atmospheric pressure p ₁ ;

ν is the exponent in the law of the rate of combustion of TRT;

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

Γ (k) is the function of the adiabatic exponent k defined by equation [7]

.

.

Substituting (3), (4) into equation (1), we obtain

The value of the critical cross section of the solid propellant nozzle S _cr is determined from the Bori equation [7] for the solid propellant rocket propulsion mode with the value of the working pressure in the combustion chamber p ₀ :

From the Bori equation (6) it follows:

Substituting (7) in equation (5), we obtain:

We introduce the parameter n equal to the pressure ratio

.

.

Replacing in the right-hand side of (8) the pressure p _k through p ₀ taking into account this parameter (p _k = n p ₀ ), we obtain:

In view of (10), the condition for quenching the solid fuel charge when depressurizing (2) will take the form:

The condition of non-extinction of the charge during depressurization corresponds to the inequality opposite (2):

Taking into account (10), (11), it is possible to obtain the value of the minimum free volume of the combustion chamber V _min at which the TPT charge is not extinguished when the pressure is released:

where parameter A includes the characteristics of the solid fuel charge

.

.

In accordance with (12), the minimum value of V _min is determined by the characteristics of solid fuel (parameter A), the type of fuel is mixed or ballistic (parameter B), the pressure in the combustion chamber is p ₀ during the solid propellant mode of operation of the solid propellant, and the opening pressure of the nozzle plug is p _k = np ₀ .

Implementation examples

We will calculate the minimum value of the free volume of the solid propellant combustion chamber V _min when using the end charge of ballistic (gunpowder N) and mixed (CYN) solid fuels in an active-reactive projectile with a caliber of 150 mm (S _t = 177 cm ² ) at a value of p ₀ = 4 MPa . The characteristics of these fuels are given in table 1.

The calculation results according to equation (12) of the minimum value of the volume of the combustion chamber for these fuels are shown in FIG. 3 for different values of the ratio n = p _to / p ₀ . With increasing pressure p _k (or, equivalently, parameter n), a large value of the free volume of the combustion chamber V _min is required, which ensures stable combustion of the charge when the pressure is released. The optimal value of p _k depends on the burning stability characteristics of a particular solid fuel composition and is determined, as a rule, experimentally.

The results of similar calculations carried out for the considered fuels at different values of the operating pressure p ₀ in the combustion chamber at the main operation of the engine are shown in FIG. 4. From the graphs (Fig. 4) it follows that with an increase in the operating pressure p ₀ from 4 to 12 MPa, the required value of the free volume of the combustion chamber decreases.

The calculation results shown in FIG. 3, 4, show that when using mixed solid fuels, a significantly lower value of V _min is required than for ballistic compositions.

Thus, the inventive rocket engine of an active-rocket projectile ensures the achievement of the technical result of the invention - improving the reliability of the initiation and operation of the mid-flight rocket engine due to an autonomous igniter, eliminating the effects of high-pressure propellant gases in the gun barrel on the rocket engine, and also by ensuring combustion stability the charge of the rocket engine when opening the nozzle plug after the projectile leaves the barrel.

Information sources

1. Serebryakov M.E. Internal ballistics. M .: Oborongiz, 1949 .-- 670 p.

2. RF patent №2021544, IPC F02K 9/08. The rocket engine of an artillery shell / Sokolov G.F., Mironov Yu.I., Berkovich B.C. - Publ. 10/15/1994 g.

3. RF patent No. 2037065, IPC F02K 9/08. Rocket engine of solid propellant rocket / Sokolov G.F., Vasina E.A., Morozov V.D., Koshelev E.V. - Publ. 06/09/1995 g.

4. RF patent No. 2059859, IPC F02K 9/08. Rocket engine of an artillery shell / Babichev V.I., Aleshichev I.A., Sokolov G.F., Rodin L.A. - Publ. 05/10/1996

5. RF patent No. 2075033, IPC F42B 10/38. Active rocket / Babichev V.I., Kolotilin S.V. - Publ. 03/10/1997

6. RF patent No. 2080468, IPC F02K 9/08, F42B 10/38. Rocket engine of an artillery shell / Babichev V.I., Sokolov G.F., Mironov Yu.I., Berkovich B.C. - Publ. 05/15/1997

7. Reisberg B.A. Fundamentals of the theory of working processes in solid fuel rocket systems / B.A. Reisberg, B.T. Erokhin, K.P. Samsonov. - M.: Mechanical Engineering, 1972. - 383 p.

8. Prisnyakov V.F. Dynamics of solid fuel rocket engines. Textbook for high schools. - M.: Mechanical Engineering, 1984. - 248 p.

9. Shishkov A.A. Workflows in solid propellant rocket engines: Directory / A.A. Shishkov, S.D. Panin, B.V. Rumyantsev. - M.: Mechanical Engineering, 1988 .-- 240 p.

Claims (17)

A rocket engine of an active-reactive projectile containing a combustion chamber with a charge of solid fuel, a nozzle, a pyrotechnic initiator and a nozzle plug, characterized in that a breakthrough membrane is installed in the critical section of the nozzle, the plug consists of a base, a cover and a perforated hollow cylindrical glass fixed to the base bottom side of the membrane, at the base of the plug and the bottom of the glass are made coaxial holes in which the rod is mounted with the possibility of longitudinal movement, the rod has a pointed the end on the membrane side, a conical thickening on the side of the base of the plug, mating with a conical recess in the base, and a cut flange sandwiched between the base and the cover of the plug, a console is fixed on the rod inside the glass, a cylindrical spring is installed in the glass between its bottom and the console, covering the rod moreover, the pyrotechnic initiator consists of a sample of the main igniter, placed between the bottom of the glass and the membrane, and at least two igniter caps mounted on the base of the plug and with drummers mounted on the console, while the cap of the nozzle plug is located in the outlet section of the nozzle and secured by rolling from its outer side, a hole is made in the central part of the cap, the diameter of which is equal to the diameter of the conical thickening of the rod, and the free volume of the combustion chamber between the charge solid fuel and the nozzle inlet section corresponds to the inequality

, where V is the free volume of the combustion chamber, m ³ ; A = kƒ _p ρ _T S _T u ₁ ; k is the adiabatic exponent of the solid fuel combustion products; ƒ _p = RT _p is the reduced power of the fuel, J / kg; R is the gas constant of the solid fuel combustion products, J / (kg⋅K); T _p - adiabatic combustion temperature of solid fuel at constant pressure, K; ρ _t - density of solid fuel, kg / m ³ ; S _t - surface area of the combustion of a charge of solid fuel, m ² ; u ₁ - burning rate of solid fuel at atmospheric pressure, m / s; p _{to the} opening pressure of the nozzle plug, Pa; p ₁ = 0.1 MPa - atmospheric pressure; n = p _to / p ₀ - pressure ratio; p ₀ - pressure in the combustion chamber at the marching mode of operation of the rocket engine, Pa; ν is the exponent in the law of the burning rate of solid fuel;

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