RU2253083C1 - Propelled component - Google Patents

Abstract

FIELD: propelling equipment, applicable, in particular, for enhancing the efficiency of propelling devices and jet technology.

SUBSTANCE: the component has a massive central subcaliber-type core in the form of an extended cylinder with a sharpened front part and with tail fins in the rear part, leading device-obturator consisting of radially joined sections with the core of the propelled component fastened in its inner hole. The front end face surface of the obturator is made with a ring-shaped recess, concentric relative to the common axis of the core, leading device. Positioned above the outer rear part of the obturator is the casing of the propelled component, whose central axis coincides with the common axis of the core and obturator, the rear part of the mentioned casing rests on the ends of the fins, solid propellant is positioned in the space of the combustion chamber confined by the inner side surface of the casing outer surface of the core, rear end face surface of the obturator and the rear plane of the casing. The outlet section of the combustion chamber is divided into ring sections by the core fins, each section of the obturator has at least one through duct connecting the ring-shaped recess of the front end face surface of the obturator to its end face rear surface, the mentioned duct and its continuation in the space of the combustion chamber are filled with quick-burning propellant, and the rest section of the combustion chamber is filled with slow-burning propellant.

EFFECT: the following characteristics are improved: the absolute value of the on-board power reserve, useful power, power supply in the boost stage, economical efficiency, reliability, mass and dimensional characteristics, cost.

6 cl, 4 dwg

Description

The invention relates to military equipment and can be used, in particular, to increase the effectiveness of armor-piercing artillery and rocket technology.

A known analogue of the proposed one is “Cartridge with reactive penetrating part” [1] (Patent RU 2099667 Int.Cl. F 42 B 5/02, priority 05.27.94), intended for ballistic dispersal in the barrel of a propelling device (MU) and for marching reactive do-disperse a missile element (ME) containing a massive central core with a pointed frontal part, a casing is placed over the outer stern of the missile element, the central axis of which coincides with the axis of the core, the volume of the combustion chamber limited by the inner side surface of the housing and the feed plane of the body, filled with volumetric separated parts of fast-burning and slowly burning solid fuels, which coincides with the essential features of the proposed method.

In this case, slowly burning fuel (of the first type) is initiated at the initial moment of the launch of the projectile, and fast-burning (of the second type) after it leaves the barrel of the propelling device.

The disadvantage of analogue [1] is:

- using only its on-board fuel supply for marching overclocking of MEs and in a corresponding reduction in the kinetic energy of MEs at the end of overclocking, necessary to accomplish the task;

- increased air resistance during flight and a corresponding decrease in the kinetic energy of the ME, the range of its flight, which is due to the large-caliber core ME.

Closest to the proposed technical solution is adopted as a prototype, “The core of a sub-caliber rocket projectile [2] (Pat. US No. 4703696 N.C1.102 / 517 Pr.03.11.1987), intended for acceleration in the barrel MU, containing a massive central core in the form of an extended cylinder with a pointed frontal part and with tail feathering in the aft part, a sealer consisting of n radially articulated sections, through which an ME core is threaded, the frontal end surface of the obturator is made with an annular angle ubleniem centered on a common axis of the core, the obturator and the housing, which coincides with the essential features of the proposed method.

In this case, the fuel of the chamber charge of the propelling device (expelling charge) is made quick-burning, its ignition is initiated at the initial moment of the launch of the projectile.

The operation of the prototype device is based on the following principles.

It is known that increasing the firing range requires an increase in the acceleration rate of the ME in the MU barrel and a decrease in its aerodynamic drag at the stage of marching flight. These requirements are contradictory, since it is advisable to increase the bottom area of the ME (to increase the force pushing the ME out of the barrel) to increase the muzzle velocity of the MOH, however, the air flow resistance, which slows down the ME with the shutter on the march, increases according to Stokes’s law. Taking into account the noted ME, it is made in the form of a massive sub-caliber type core in the form of an extended cylinder with a pointed frontal part and with a tail in the aft part, to which a leading element (obturator) is attached, which corresponds to the caliber of the MU barrel. The shutter consists of radially articulated sections for the convenience of their subsequent separation from the core.

Accordingly, in the prototype, a missile element designed for combined dispersal (in the barrel of a throwing device and mid-flight reactive dorazgon) is placed in the barrel of a throwing device. At the moment of ignition, the fuel begins to produce gases, the pressure in the barrel increases and overcomes the adhesion force of the cartridge case to the side surface of the obturator, knocks the obturator out of the specified body, and then pushes the obturator together with the core attached to it from the barrel of the throwing device. Due to the presence of an annular groove on the frontal surface of the obturator, a radial pressure force of the incoming air flow is formed, directed from the center and disconnecting the obturator sections outside the barrel. After that, the core of small cross section continues to fly without a shutter. The entire energy reserve of the spent fuel is further distributed between the costs of the work against the resistance force of the atmosphere and the kinetic energy of the ME itself. The increase in this energy by increasing the mass of the chamber charge is associated with a significant deterioration in the operational characteristics of the ME (an increase in the mass and size parameters of the MU - barrel length, wall thickness, mass). Given the preferred limitation of these characteristics, the required kinetic energy of armor-piercing ME in modern MU is not achieved.

This is the main disadvantage of the prototype method. In addition, with a limitation of the achieved speed, the time of free ballistic flight to the target is too long, which increases the spread of the realized flight parameters, reduces accuracy, and the aiming range.

All this leads to a decrease in the efficiency of using ME (including the efficiency of the installation), and a decrease in the specific impulse and kinetic energy of ME realized during acceleration, which is also a significant disadvantage of the prototype. In addition, when trying to increase the kinetic energy of ME due to ballistic barrel acceleration, the recoil momentum of MI is inevitably increased, which is also a serious drawback of traditional throwing equipment.

So, the disadvantage of the prototype device [2] is the deterioration of the following characteristics:

- the absolute magnitude of the power ratio (kinetic energy), including due to the lack of energy supply on the marching section of the trajectory;

- Efficiency (COP), as well as

- weight and size characteristics of the complex (ME + MU) and

- the reliability and service life of a regular MU due to the increased recoil momentum, as well as due to increased temperature and pressure in the entire volume of the internal cavity of the barrel and, accordingly,

- cost.

Accordingly, the technical result required when implementing the device consists in eliminating the above disadvantages.

List of drawing figures.

Figure 1 - diagram (longitudinal section) of the proposed device;

Figure 2 - diagram (rear view) of the proposed device;

Figure 3 - diagram (rear view) of the proposed device according to claim 3 of the formula;

Figure 4 - diagram (top view) of the articulated mechanism of the keels of the ME with the deployed (a) and folded (b) positions of the keels (according to claim 4, 5 of the formula).

In the drawing of figures 1-4, the following conventions of the constituent elements are used:

1 - ME core;

2 - obturator (leading element) ME;

3 - casing ME;

4 - keel of the tail of the ME;

5 - fast burning solid fuel;

6 - fast burning solid fuel;

7 - frontal depression of the obturator;

8 - channel obturator ME;

9 - keel hinge;

10 - keel spring;

11 - sector of the output section of the combustion chamber;

12 - lateral side of the keel;

13 - aft of the lateral side of the keel;

14 - fixed part of the side of the keel.

Figure 1 presents a diagram (longitudinal section) of the proposed device, which shows the relative position of the structural elements of the ME.

Figure 2 presents a diagram (rear view) of the proposed device, which shows the location of the aft edges of the keels ME;

Figure 3 presents a diagram (rear view) of the proposed device, which shows the open arrangement of the movable elements of the keels with the placement of slowly burning fuel in the internal volume of the keels.

Figure 4 presents a diagram (top view) of the hinge mechanism of the keels of the ME with the deployed (A) and folded (B) positions of the keels, which shows the location of the springs and hinges connecting the movable elements of the keels with fixed ones.

To eliminate the disadvantages of the prototype method, a missile element is proposed that contains a massive central calibrating-type core in the form of an extended cylinder with a pointed frontal part and tail in the aft, a shutter master device consisting of radially articulated sections, in the inner hole of which a missile element core is fixed , over the outer stern of the obturator is placed a cover of a missile element, the central axis of which coincides with the common axis of the core and obturator, core The new part of the specified casing rests on the ends of the tail fin keels, and in the volume of the combustion chamber bounded by the inner side surface of the casing, the outer surface of the core, the aft end surface of the shutter and the aft plane of the casing, solid fuel is placed, the output section of the combustion chamber is divided into ring sectors by the tail keels the feathering of the core, and the frontal end surface of the obturator is made with an annular recess concentric with respect to the common axis of the core, leading about element and casing

Moreover, each of the sections of the obturator has at least one through channel connecting the annular recess of the frontal end surface of the obturator with its end aft surface, this channel and its continuation in the volume of the combustion chamber are filled with quick-burning fuel, and the rest of the combustion chamber is filled with slow-burning fuel .

In addition, the sides of the tail fin carcasses are made converging at an acute angle directed at its apex to the frontal part of the core, and the volume of the combustion chamber bounded by the sides of each keel, the outer surface of the core and the inner surface of the body is filled with slow-burning fuel.

In addition, the sides of the tail fin keels are made up of front fixed and rear movable sections connected by hinges, the axes of which lie in a plane passing through the central axis of the missile element, and the sides of each keel are interconnected by springs.

In addition, the solid fuel filling the longitudinal channel of each of the shutters and its continuation in the volume of the combustion chamber, as well as the fuel filling the rest of the combustion chamber, are ignited simultaneously with the initiation of the chamber charge of the propelling device.

In addition, spacer springs connecting the lateral sides of each keel are made with the possibility of collapse and fixation in this state when their compressing force of the incoming air flow reaches a threshold level.

In addition, the casing of the missile element and the radially articulated sections of the shutter are made detachable from each other and from the core after the combustion of slowly burning fuel.

Consider the work of the proposed device, made according to the scheme of figure 1. A missile element designed for combined dispersal (in the barrel of a throwing device (MU) and a marching rocket-direct-flow jet dorazgon) is placed in the barrel of the throwing device. The leading element, which takes on the gas pressure in the MU barrel, is the obturator 2, consisting of radially articulated sections, through the inner opening of which the core 1 of the missile element is threaded. On top of the outer aft part of the obturator there is a housing (casing) 3 of a missile element, the central axis of which coincides with the common axis of the core 1 and the obturator 2. The aft part of the specified casing 3 is based on the ends of the tail fin 4 to impart sufficient strength to the structure subjected to high acceleration accelerations. The volume limited by the inner side surface of the housing 3, the outer surface of the core 1, the aft end surface of the shutter 2 and the aft plane of the housing 3 forms a combustion chamber, it is filled with solid quick-burning (5) and slow-burning (6) fuel. The open aft end of the casing 3 creates an output section of the combustion chamber. It is divided into annular sectors by keels 4 of the tail unit of core 1. In each of the annular sectors of the output section of the combustion chamber, a channel 8 is formed, filled with quick-burning solid (or pasty) fuel 5. These channels 8 extend to the frontal end surface of the seal 2, passing slowly through the thickness burning solid fuel 6, and then fills the channels in the thickness of the shutter itself 2. The fuel of the first 5 and second 6 types are initiated simultaneously with the aft of the combustion chamber. The channels of quick-burning fuel 8 have time to burn out at the stage of barrel acceleration of MEs along with gasification of the MU chamber charge. On a ballistic section of the free flight trajectory, the incoming air flow strikes into the annular recess 7 of the frontal end surface of the obturator 2, focuses it, passing in a compressed state into the through channel 8 of the obturator, which thus performs the functions of the air intake of an air rocket-direct-flow engine (RPD). The power ratio of such an engine is greater than that of a standard solid-fuel engine of the prototype, since the energy and mass of atmospheric air is additionally used.

After the burnup of fuel 6 and the casing 3, which ensured the operation of the RPD, the pressure of the incoming air flow directed from the center separates the sections of the obturator 2 outside the barrel, minimizing the drag of the ME at the final stage of the ballistic trajectory.

As noted above, the output section of the combustion chamber is divided into sectors using keels 4 of the tail unit (figure 2). To clarify the design of such a keel 4, imagine it consisting of the left and right sides 12.

The lateral sides of the 12 keels 4 of the tail unit of the core 1 are made diverging at an acute angle directed by its apex to the frontal part of the core 1. Thus, the area of each of the sectors 11 of the output section of the combustion chamber is limited by the opposite (right and left) side 12 sides of the adjacent keels 4 ( figure 3). In this case, the cross section of the flow of gas entering from the chamber narrows in the feed nozzle between the keels inversely to the angle α formed by the lateral sides 12 of each of the keels 4. This is necessary to increase the pressure in the combustion chamber (throttling) to the optimal value. Part of the fuel 6, made with a large constant combustion time, is placed in the volume of the combustion chamber, limited by the lateral sides 12 of each of the keels 4, the outer surface of the core 1 and the inner surface of the housing 3 (figure 4).

At the stage of free ballistic flight after the burning of slow burning fuel 6, the aft parts 13 of the sides 12 of each tail fin are brought together (Fig. 4, B), which increases their aerodynamic characteristics in free flight conditions. For this, each of the sides 12 of each keel 4 is made up of a front fixed (14) and aft (movable) (13) sections connected by hinges 9, the axes of which lie in a plane passing through the central axis of the missile element 1, and the sides 12 of each keel 4 are interconnected by spacer elements 10. Spacer elements (for example, springs) provide the necessary position of the stern sections of the keels, counteracting the force of the aerodynamic pressure on them of the flow of incoming gas.

The implementation of the spacer elements 10 in the form of elastic elements (for example, springs) allows to stabilize the gas pressure in the combustion chamber. Indeed, with a decrease in pressure in the flow of the gas leaving the nozzle, the springs burst the feed elements 13 of the keels 4, which leads to the compression of the specified flow and increase the pressure in the combustion chamber.

It is clear that the lateral sides 12 of each keel are under the influence of mutually opposite pressure forces:

- from the side of the gas stream flowing from the sector of the output section of the combustion chamber (forces compressing the spring), and

- from the side of the gas stream flowing from the corner formed by the dilated lateral sides of the keel (forces stretching the spring).

It should also be noted that the additional compression of the gas stream flowing out of the sector of the output section of the combustion chamber is carried out by jets of gas leaving the intrafeature space. As the fuel 6 burns out, the pressure of the flow of burning gas decreases and the main force on the sides 12 of the keels 4 exerts pressure of the atmospheric flow. At some point, the resulting bursting force of the springs 10 overcomes a predetermined threshold value. Upon reaching the threshold force compressing the spacer springs 10, they collapse and lock in this state, which improves the aerodynamics of the ME free flight, as indicated above.

The casing of the projectile element and the radially articulated sections of the shutter are separated from each other and from the core after the combustion of slowly burning fuel, are thrown to the side by the forces of the incoming flow resistance, improving the aerodynamics of the ME free flight.

As already noted, solid fuel 5, filling the longitudinal channel of each of the shutters and its continuation in the volume of the combustion chamber, as well as fuel 6, filling the rest of the combustion chamber, is ignited simultaneously with the initiation of the chamber charge of the propelling device. This increases the initial momentum of the ME, informs it of the additional speed at the stage of acceleration in the trunk of the MU.

The expediency of the simultaneous combustion of all types of fuel in the barrel is explained by the fact that increasing the speed due to an increase in pressure in the ME combustion chamber at the stage of muzzle acceleration does not require thickening of its walls, since the surface of the barrel channel takes up radial loads. Otherwise, an increase in the mass of the combustion chamber, and, therefore, of the entire ME, would decrease the acceleration of the ME at the stage of marching overclocking in accordance with Newton’s second law. Thus, there is no need to provide increased mechanical strength and temperature stability of the combustion chamber of the ME, which significantly reduces its weight and size parameters, part of the resource of which can be spent on power output (mass of core and fuel), thereby increasing kinetic energy, accuracy and range .

Next, we show that it is thanks to the significant differences of the proposed method that the required technical result is provided.

The fact that each of the sections of the obturator has at least one through channel connecting the annular recess of the frontal end surface of the obturator with its end aft surface, this channel and its continuation in the volume of the combustion chamber are filled with solid fuel of the first type with a small combustion time constant and the fuel of the second type, filling the rest of the combustion chamber, is made with a large constant combustion time, provides:

- increasing the absolute value of the energy reserve (kinetic energy) and efficiency (efficiency), range, accuracy, reliability of the booster due to reduced recoil, as well as due to increased temperature and pressure in the entire volume of the internal cavity of the barrel, as well as

- reduction of weight and size characteristics and cost, taking into account the following:

- the energy consumption of the core for aerodynamic drag is not only offset on the marching portion of the trajectory, but also provides an increase in the kinetic energy of the core due to combustion of fuel;

- the energy supply of not only on-board fuel, but also outboard atmospheric oxygen is realized;

- the implementation of these energy reserves is carried out in an optimal way, taking into account the launch of the ramjet mechanism on the marching section of the ME flight path;

- the possibility of increasing the energy of barrel acceleration without increasing the weight and size of the combustion chamber of the ME;

- the energy of the fuel supply is realized gradually not only at the stage of muzzle acceleration, but also on the marching section of the flight path.

The fact that the solid fuel filling the longitudinal channel of each of the shutters and its continuation in the volume of the combustion chamber, as well as the fuel filling the rest of the combustion chamber, is ignited simultaneously with the initiation of the chamber charge of the propelling device, increases the core energy increment at the muzzle acceleration stage, increases reliability the operation of the ignition mechanism of various types of fuel (and the device as a whole), simplifies the design of the device, reduces its cost;

The fact that the sides of the keels of the tail feathers of the core are made diverging at an acute angle directed at the apex to the frontal part of the core, and the volume of the combustion chamber bounded by the sides of each keel, the outer surface of the core and the inner surface of the body is filled with fuel made with a large time constant combustion, additionally increases the energy resource of the device, optimizes the ramjet operation mode, therefore, increases the overall efficiency of the device;

The fact that the sides of the tail fin keels are made up of front fixed and rear movable sections connected by hinges, the axes of which lie in a plane passing through the central axis of the missile element, and the sides of each keel are connected by spacer springs, allows you to switch to the optimal aerodynamic configuration of the device during the transition to inertial ballistic flight;

The fact that the springs connecting the lateral sides of each keel are made with the possibility of collapse and fixation in this state when their compressive force reaches a threshold level, also allows you to switch to the optimal aerodynamic configuration of the device when switching to inertial ballistic flight;

The fact that the casing of the missile element and the radially articulated sections of the shutter are made detachable from each other and from the core after the combustion of slow-burning fuel also allows you to switch to the optimal aerodynamic configuration of the device when switching to inertial ballistic flight.

At the same time, it is possible to use the resulting effect to simultaneously improve the overall, maneuverable, technical, economic and power characteristics of various types of equipment for fuel, energy and defense complexes. For example, part of the gain in the energy resource can be directed to the simultaneous improvement of a number of parameters: reduction in size, cost and increase in the range and reliability of the device.

Thus, it is shown that the required technical result is really achieved due to the significant differences of the proposed installation.

The experiments showed the feasibility of the invention.

Claims (6)

1. A missile element containing a massive central subcaliber type core in the form of an extended cylinder with a pointed frontal part and a tail in the aft, the leading device is a sealer consisting of radially articulated sections, in the inner hole of which a missile element core is fixed, frontal end surface the shutter is made with an annular recess concentric with respect to the common axis of the core and the master device, characterized in that on top of the outer stern of of the tyurator there is a cover of the projectile element, the central axis of which coincides with the common axis of the core and the shutter, the aft of the specified cover is supported on the ends of the tail fin keels, and in the volume of the combustion chamber bounded by the inner side surface of the casing, the outer surface of the core, the aft end surface of the shutter and the aft plane of the casing, solid fuel is placed, the output section of the combustion chamber is divided into annular sectors by the keels of the tail unit of the core, each of the sections The ator has at least one through channel connecting the annular recess of the frontal end surface of the obturator with its end aft surface, the specified channel and its continuation in the volume of the combustion chamber are filled with quick-burning fuel, and the rest of the combustion chamber is filled with slow-burning fuel. 2. The element according to claim 1, characterized in that the charges of the slow and fast burning fuel are configured to initiate simultaneously with the initiation of the chamber charge of the propelling device. 3. The element according to claim 1, characterized in that the frontal parts of the side walls of the keels of the tail of the core are converging at an acute angle directed at its apex to the frontal part of the core, the inner volume of each of the keels limited by these side walls, the outer surface of the core and the inner surface of the casing, filled with slow-burning fuel. 4. The element according to claim 3, characterized in that the lateral sides of the tail fin keels are made up of front fixed and rear movable sections connected by hinges, the axes of which lie in a plane passing through the central axis of the missile element, and the lateral rear sides of each keel interconnected by springs. 5. The element according to claim 4, characterized in that the spacer springs connecting the lateral rear sides of each keel are made with the possibility of collapse and fixation in this state, when the compressive pressure force of the incoming air stream reaches a threshold level. 6. The element according to claim 4, characterized in that the casing of the missile element and the radially articulated sections of the obturator are made detachable from each other and from the core after the combustion of slow burning fuel.

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